42349861 Design of Liquid Propellant Engines Textbook

d

_

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_/

:_'_

_'

,-7{ "

_';""_"'-7 ;, _, 77:7>

i

.

/

i

I

-I

Declasslficd

i

by _ut._or!ty

df NA_A-,,._

i

i I

.!

!ii i :7 d

P=

of e PO

Optimum thrust for a given ambient pressure is obtained when the nozzle expansion area ratio e=Ae/At is such that Pe--Pa. This may be seen from figure 1-10. If the divergent nozzle section is extended in the region where Pe > Pa, thrust will increase. Where Pe

uJ 5200 I0

20

30

40

50

CHAMBER CHARACTERISTICLENGTH( Le) IN. Figure

4-7.-Effect 6[ L* on c* value mental thrust chamber.

4-1.-Recommended

Characteristic Length lant Combinations

of experi-

Combustion (L*) for Various

Propellant combination

Chlorine trifluoride/hydrazine-base fuel.. Liquid fluorine/hydrazine ............. Liquid fluorine/liquid hydrogen (GH_ injection) .......................... Liquid fluorine/liquid hydrogen (LH 2 injection_ .......................... Hydrogenperoxide/RP-I (including catalyst bed) ....................... Nitric acid/hydrazine-base fuel ........ Nitrogen tetroxide/hydrazine-hase fuel.. Liquid oxygen/ammonia ............... Liquid oxygen/liquid hydrogen (GH2 injection) .......................... Liquid oxygen/liquid hydrogen (LH 2 injection) .......................... Liquid oxygen/RP-1 ..................

Combustion in

various thrust chamber designs. Typical L* values for different propellants are given in table 4-1. With At and minimum required L* established, the required combustion chamber volume Vc can be calculated by equation (4-4).

NZ041 50-50 O/F MIXTURE

TABLE

Chamber Propel-

Combustion chamber characteristic length (L*), m. 30-35 24-28 22-26 25-30 60-70 30-35 30-35 30-40 22-28 30-40 4@50

Chamber Shape

As can be seen from equation (4-3), the stay time ts is independent c; the combustion chamber geometry. Theoretic:ally, for a given required volume, the chamber can be of any shape. In actual design, however, the choice of the combustion chamber configuration is limited. In a long chamber with a small cross section, higher nonisentropic gas flow pressure losses will result as explained in chapter I. This approach also dictates a longer thrust chamber space envelope and imposes certain space limitation on the injector design to accommodate the necessary number of injector holes. With a short chamber of large cross section, the propellant atomization or vaporization zone occupies a relatively large portion of the chamber volume, while the mixing and combustion zone becomes too short for efficient combustion. Other factors, such as heat transfer, combustion stability, weight, and ease of manufacturing, are to be considered in determining the final combustion chamber configuration. Three geometrical shapes which have been used m combustion chamber design are shown in figure 4-8. While the spherical and the nearspherical chambers were used in earlier European

v

88

v

__

--

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

designs, the cylindrical chamber has been used most frequently in the United States. The spherical or nearly-spherical chamber, as compared to the cylindrical one of the same volume, offers the advantage of less cooling surface and weight. A sphere has the smallest surfaceto-volume ratio. For equal strength of material and chamber pressure, the structural walls of the spherical chamber are about half the thickness of the walls of a cylindrical chamber. However, the spherical chamber is more difficult to manufacture and has poorer performance under most circumstances. For these practical reasons, the design details of the cylindrical combustion chamber will be treated in this book. Several

THROAT

/NOZZLE

novel thrust chamber designs will also be discussed. In the design layout of the cylindrical combustion chamber of a given At and L*, the value of the contraction area ratio, (ec =(Ac/At)) can be optimized through careful studies of the following factors: (1) Combustion performance in conjunction with the injector design (2) Chamber gas flow pressure drop (3) Chamber wall cooling requirements (4) Combustion stability (5) Weight (6) Space envelope (7) Ease of manufacturing For pressurized-gas propellant feed, lowthrust en_ne systems contraction area ratio values of 2 to 5 have been used. For most turbo-

pump propellant feed, high thrust and high chamengine systems lower ratio values THRUST--CHAMBER ber pressure of 1.3 to 2.5 are employed. The reader is also referred to section 1.2 chapter I, "The Gas-flow Processes in the Combustion Chamber and the AXIS Nozzle."

IN_

The basic elements of a cylindrical combustion chamber are identified in figure 4-9. In design practice, it has been arbitrarily defined that the combustion chamber volume includes the

COMBUSTION CHAMBER

IN_

space between injector face I-I and the nozzle throat plane II-II. The approximate value of the combustion chamber volume can be expressed by the following equation

NOZZLE

Vc = A_ [Lcec + ½_-A-_Cot _-NEAR SPHERICAL COMBUSTION CHAMBER

INJECTOR

(4-5)

THROAT

FACE

[-CHAMBER

I

T_T

O(ec_/3 -1)]

NOZZLE

[

DIA

I

AREA

"

n" Dc

I

Ac

,

,---

NOZZLE

/

INJECTOR

?

HRUSTC.___.HHAMBER CHAMBE:"

AXIS [

[

_

CYLINDRICAL

OIA

[

AREA

Dt A t

SECTION

z_LENGTH

CYLINDRICAL COMBUSTION CHAMBER Figure

4-8.-Frequently used geometrical [or combustion chambers.

Lc

CHAMBER

CONTRACTION AREA

shapes

Figure

RATIO

Ac (¢

= A'-'_T

4-9.-Elements of basic cylindrical bustion chamber.

com-

DESIGNOF THRUST CHAMBERSAND OTHER COMBUSTIONDEVICES

The total

surface

area of the combustion

chamber walls excluding injector face can be approximated by the following expression:

Total

Nozzle

area =2Lc_cAt

Expansion

It was learned

+ csc 9(ec - 1)At (4-6)

Area Ratio earlier

that

with all other

parameters fixed, in particular chamber pressure, there is only one optimum nozzle expansion area ratio for a given altitude or, more specifically, ambient pressure. Except for those systems which start in vacuum, ambient pressure will have to be considered. This is especially true for boosters which start at or near sea-level conditions. It is the ultimate purpose of a rocket engine to lift vehicles to altitudes. Inherently, then, ambient pressure will not be a constant (except for high-altitude starts, as mentioned). It is, therefore, extremely important for the designer to know the trajectory of the vehicle to be propelled or, more specifically, its altitude-versus-time characteristics. With this information, the designer is in a position to make a first, optimizing selection of a nozzle expansion area ratio, for best results throughout the entire trajectory. As shown earlier, area ratio will be truly optimum for only one specific altitude. The optimization for ambient pressure then is essentially an averaging process. Other considerations usually cause the designer to deviate from the "paper optimum" for the nozzle expansion area ratio. Some of the most common are: weight, size, ease of manufacturing, handling, and cooling (heat transfer) considerations.

Nozzle

89

The selection of an optimum nozzle shape for a given expansion area ratio is generally influenced by the following design considerations and goals: (1) Uniform parallel axial gas flow at the nozzle exit for maximum momentum vector (2) Minimum separation and turbulence losses within the nozzle (3) Shortest possible nozzle length for minimum space envelope, weight, wall friction losses, and cooling requirements (4) Ease of manufacturing In actual design practice, any abrupt change or discontinuity in the nozzle wall contour should be avoided to eliminate the possibility of shock waves or turbulence losses. Theoretically, the nozzle throat is simply the unique plane of minimum cross-section area. In practice, a wellrounded throat section is employed. Only at the nozzle exit plane is a sharp edge used because a rounded one would permit overexpansion and flow separation. 1. Conical

Nozzle

In early rocket engine applications, the conical nozzle, which had proved satisfactory in most respects, was used almost exclusively. The advantages of a conical nozzle are ease of manufacturing and flexibility of converting an existing design to higher or lower expansion area ratios without major redesign of the nozzle contour. The configuration of a typical conical nozzle is shown in figure 4-10. The nozzle throat section has the contour of a circular arc with a radius R ranging from 0.5 to 1.5 times the throat radius Rt. The half angle of the nozzle convergent cone section can range from 20 ° to 45 °. The

Shape

Most rocket nozzles are of the convergingdiverging De Laval type. Since the flow velocity of the gases in the converging section of rocket nozzle is relatively low, any smooth and wellrounded convergent nozzle section will have very low energy losses. By contrast, the contour of the diverging nozzle section is very important to performance, because of the very high flow velocities involved.

RI

,. Figure

:4-I O.-Conical

y-os nozzle

contour.

9O

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

divergent cone half angle a varies from approximately 12 ° to 18 °. The length of the conical nozzle section can be expressed by the equation

Rt(_e - 1) + R(sec Ln= tan a

/ / /

E.

L_E

_-c:,,-_,.

1

(4-7)

-,---__

;_= _ (1 + cos a)

(4-8)

where a = half angle of the conical nozzle. For an ideal nozzle, A would be unity. For a conical nozzle with a = 15 ° and h =0.983, the exit gas momentum or the exit velocity will be 98.3 percent of the ideal nozzle exit velocity calculated by equation (1-18). The value of the vacuum thrust coefficient of a nozzle is in direct proportion to the thrust generated by the nozzle, or to the nozzle exit gas velocity. Therefore, the theoretical vacuum thrust coefficient (neglecting friction and other flow losses) of a conical nozzle with 15 ° half angle will be 98.3 percent of the ideal nozzle thrust coefficient calculated by equation (1-33a). 2. Bell Nozzle For increased performance and shorter length, bell-shaped nozzles have been developed. This nozzle design employs a fast expansion or radial flow section in the initial divergent region, which then leads over to a uniform, axially directed flow at the nozzle exit. The wall con-

zle.

.

/

with uniformly parallel axial gas flow. The value of h can be expressed by the following equation:

enough

PLANE L_

[:

a - 1)

The conical nozzle with a 15 ° divergent half angle has become almost a standard, as it is a good compromise on the basis of weight, length, and performance. Since in a conical nozzle certain performance losses occur as a result of the nonaxial component of the exhaust gas velocity, a correction factor h is applied for the calculation of the exit gas momentum. This factor or thrust efficiency is the ratio between the exit gas momentum of the conical nozzle and that of an ideal nozzle

tour is changed gradually shocks will not form.

EXIT [

so that

oblique

Figure 4-11 shows the contour of a bell nozA circular arc of selected radius R, is

/ _/ _,

1 Figure

....., C_CrERIS_r

ic

._ 4-11.-Bell

nozzle contour.

chosen for the nozzle contour MT upstream of the throat. Contour TNE is the diverging portion of the nozzle. The initial expansion occurs along contour TN; contour NE turns the flow over to a direction nearer to axial. For design convenience, the contour TN is also a circular arc, with a smaller radius R2. For those familiar with compressible flow theories, it is noted that, using transonic flow analyses, a constant-Mach-number line TO can be defined at the throat. Given the flow condition along TO and the solid boundary TN, a kernel flow field TNKO can be generated by the method of characteristics developed in gas dynamics. The kernel of the rocket nozzle contour is defined as that portion of the supersonic flow field determined entirely by throat conditions. The last right characteristic line NK of kernel TNKO, and thus the location of the point N along contour TN, is to be determined by specific design criteria. The location of the end point E along contour NE is defined by the given nozzle expansion area ratio and nozzle length (distance between throat and exit plane). Then the right characteristic line NK can be determined by satisfying the following conditions concurrently: (1) A control surface PE can be generated between the point E and a selected point P along the line NK (2) Mass flow across PE equals the mass flow across NP (3) Maximum thrust by the nozzle is attained. By selecting points P', P", etc., along line NK, a series of control surfaces P'E', P"E", etc., can be generated to define points E I, E", etc., along the contour NE. Calculations for the nozzle contour can be effectively performed by a computer.

91

DESIGN OF THRUST CHAMBERS AND OTHER COMBUSTION DEVICES

8,\

I

I

IO0

99

!

.J

j

-_ 98 pZ _J {J nr bJ O.

/

!

,jo POINT

NOZZLE AX*S

i

////

i I,

Ln

THaT 97

Figure

section

/

9_

with

T to the 60

70

90

80

I_

Lf

exit

%

NOZZLE CONICAL

Figure

LENGTH NOZZLE

4-12.-Thrust

length.

BASED ANY

ON

for

conical

ANGLE

RATIO

versus

comparison:

nozzle,

A 15 = HALF

AREA

efficiency

(Shown

ening

(Lf)

WITH

bell

effect

increasing

lowing

nozzle

Rt from

a parabola

from

the

design

data

are

Commonly, nozzle

nozzles.

used

For

percent exit

an equivalent is

bell

plane)

Throat Axial

length

of the

exit

plane,

Ln,

diameter,

fractional

angle.)

area,

nozzle

the

the

of an

between or 0.8

nozzle

below

to specify

length

(distance

80 percent,

conical

radius

15 ° half-angle

a standard

instance,

is

half-angle

as

having

throat,

the

and

(3)

Expansion

bell

(4)

Initial

(5)

and

of a 15 °

area

shows the thrust nozzle length

nozzle,

the

fol-

Dr, inches nozzle

inches

length area

wall

throat

Nozzle

exit

The

wall

angles,

4-14

as

a function

from

throat

(or the

Lf based

to

desired

on

a 15 °

ratio

angle

e

of the

parabola,

On,

wall On and of the

angle,

0e,

Oe are

degrees

shown

expansion

in figure

area

ratio

E.

expansion

ratio. Figure 4-12 versus fractional

to the

degrees

80-

same

throat

nozzle)

coni-

throat

of that

of a specific

(2)

conical

cal

the there

required:

(1)

of short-

half

of 0.382

N and

E. For

FRACTIONAL

bell

contour.

a radius

point

of

approximation

nozzle

,,,o// / /

96

4-13.-Parabolic

efficiency h LI for conical

!

40

L/= toO% Lf= 70%

and

bell As

nozzles.

may be

seen,

approximately contribute sidering

3.

nozzle

80 percent

weight

with

Approximation

convenient

nozzle

in figure upstream a radius contour

when

con-

penalties.

way

of Bell Nozzles to design

4-13.

The

of the of 1.5

Rt.

is made

2¢

a near-optimum-

is through procedures

gested by G.VR. Rao. The of a parabolic approximation diately

Lf = 80% L_= 90% Lp 100%

_5o

beyond

significantly

especially

thrust bell nozzle contour the parabolic approximation

shown

lengths

do not

to performance,

Parabolic One

bell

the

__

use of as sug-

design configuration bell nozzle is nozzle

throat The

contour

up of a circular

i

i

_ "_ 0

section entrance

EXPANSION

t

30

20

_0

imme-

T is a circular divergent

_ _'5.,c z E

AREA

40 RATIO

Lf"

60=/=

LI"

70%

Lf = 80% .f, 90°/° Lf, 100%

5O

•

arc Figure

4-14.-0n

and area

Oe as ratio

function _.

of expansion

92

DESIGN

OF

LIQUID

PROPELLANT

Optimum nozzle contours can be approximated quite accurately by selecting the proper inputs. Although no allowance is made for different propellant combinations, experience has shown that the effect of specific heat ratio y upon the contour is small. A computer program can be readily set up to perform the calculation. 4. Annular Based

theorem,

for ideal

ex-

pansion the thrust generated by a thrust chamber depends only upon the mass flow conditions (velocity and direction) at the nozzle exit. In some nozzle designs, such as annular nozzles, the gas flow at the throat is not necessarily parallel to the axis, but the exit flow is similar to that of a conical or bell nozzle and thus produces the same thrust results. There are two basic types of annular nozzles: the radial in-flow type (spike nozzle) and the radial out-flow type (expansion-deflection or E-D; reverse-flow or R-F; and horizontal-flow or H-F nozzles). They are shown in figure 4o15, together with conventional conical and bell nozAREA

ENGINES

zles. For comparison of the effect of nozzle type on size, all nozzles shown are scaled to the same thrust level, nozzle expansion area ratio, and theoretical nozzle efficiency. These nozzles show potential of adapting their geometry to space vehicle application, because shortened nozzles reduce interstage structure weight and will permit an increase in payload through increased performance for a given length. The nozzle expansion area ratio ( for an annular nozzle is defined by equation (4-9):

Nozzles on the momentum

ROCKET

Projected area of the contoured nozzle wall Ae-Ap e: Throat area At

(4-9)

where the projected area of the contoured nozzle wall equals nozzle exit plane area Ae, less the centerbody projected area Ap. Another convelent design parameter for annular nozzles is the annular diameter ratio, Dp/Dt, where Dt is the throat diameter of an equivalent circular throat, and Dp the centerbody diameter. Dp/Dt is an index of the annular

RATIO = 36:1 = 9B.3%

The parameter nozzle design

(ALT}

OF EFFICIENCY JECTOR OMBUSTION HAMBER INJECTOR COMBUSTION

HROAT I_

INJECTOR

/

rl ..--"- COMBUSTION JJ'CHAMBER

: [_ _ 'tl X

, T

:i' CONE w.,, '

Dp/D NOZZLE

LENGTH

= 100%

NOZZLE

LENGTH

= 74.2%

OVERALL

LENGTH

= I00%

OVERALL

LENGTH

= 70%

DIAMETER

= 100%

OVERALL

OVERALL

DIAMETER

_

= tO0%

"'THROAT

/

Dp/D

LENGTH

= 41.4%

OVERALL

LENGTH

= 51%

OVERALL

DIAMETER

= 105"/.

COMBUSTION

LENGTH

= 41.4%

OVERALL

LENGTH

= 519=,

OVERALL

DIAMETER

/X_

COMBUSTION

/

CHAMBER

CHAMBER {NJECTOR Ii- I

p---Dp"---'4

;_

R-F Dp/D

Y

H-F Dp/D

T = ,5

i1_

t = t0

NOZZLE

LENGTH

= 24,9

NOZZLE

LENGTH

= 14.5

OVERALL

LENGTH

= 21%

OVERALL

LENGTH

= 12 %

OVERALL

DIAMETER

= 150 %

Figure

4-I 5.-Comparison

OVERALL

of nozzle

DIAMETER

shapes.

= 194

T = 1.3

NOZZLE

INJECTOR-,,

THROAT

_

E-D ' "

I = 1.3

NOZZLE

TH R OAT

j.._D,I._,

5f 'i,!

sPIKE

,

cOM BUSTION CHAMBER

j/"-/pD'_ i

i':44--.. "i

T,'

' BELL

/

INJECTOR

i;;rl_ _../-_l.J Z"

%

%

= 102.5 %

93

DESIGN OF THRUST CHAMBERS AND OTHER COMBUSTION DEVICES

geometry The

as

compared

to a conventional

contour-calculating

nozzles

are

similar

In a conical

well

below

level

or low-altitude the

in chapter

overexpansion

low

altitudes.

characteristics,

losses.

As shown

zle

equally

(and

zles), the

the

back

the

function lower

are

4-16

back

ambient.

base

pressure

expansion the

is reached.

through

the

downstream

by

flow

the

following

(i)

The

(2)

The

two

nozzle

the

gases base

the

Figure

gas CD,

controlled

of the

are

deflected,

and

local

nozzle operation

curved-wall which

of the

turns

inner

shown

the

to some

in wall

distribution in figure

jet

gases

A typical

for low-altitude

4-16.

This

Because

corn-

CONSTANTI_ACH---k

L.NE \ TN"OAT ,NJECTO,

E

performance

AXIS BODY

wall,

as

1

4-16.-E-D

nozzle

at low altitude

that

so low

behind Since

be axial,

the

is

altitude.

of the

separation

case

inner

from

for a conven-

the

the

flow

the

gases

flow

nozzle. under

figure

4-17.

The this

ondary

spike

which

flow

secondary

streams. effect

To

describe

of truncating

is

the concept

(radial

in-

amount

nozzle

aerodynamic

flow

in

nozzle

a small

between

of

of sec-

base

spike

nozzle

geometric flow, is

the the

param-

the

manner

introduced, primary flow the

end

distri-

shown

nozzle

of secondary

energy

pressure

concept

nozzle

on

up to the

This

the

must

expansion

is also

into

of the

amount

the

spike

of various

this

relative

depending

nozzle

introduced

is a function

occur

wall

utilizes

in figure

point

nozzle.'

annular

region. Performance

the

nozzle

con-

shown

unaffected

condition

pressure

flow

closure

However,

spike

type),

as

may

may continue

bution

flow

wave

base

nozzle

at the

conditions.

of the

the

the

centerbody,

a shock

ondary

operation.

the

operation

interrelated

Figure

is

flow

becomes

in which I.

at low nature

At high-altitude verges

the

is no

which

is responsible

nozzle.

eters,

PRESSURE_

nozzle,

self-adjusting

there

is a truncated

F,,EE " SURFACE NOZZLE

WALL

altitude

wall,

nozzle

aerodynamic

'"_""_'"_

nozzle

spike

of the

nozzle

tional

at the

An improved NOZZLE WALL _

at high

for the

for improved

Pb

compression

pressure.

turning

typical

jet boundary,

influences

contour,

leads

pressure

is

also

4-17.

increases

wall

pressive

flow.

surface

nozzle operation.

the

which

Pb which

4-17.-E-D

is

boundary. Because

, Essu,

throat,

line

DE

WALL

a

initial

is

to near-axial stream

of

conditions:

contour

pressure

free

the

gases

boundary

wall

noz-

the centerbody until this

constant-Mach of the

.... '------CENTER BODY

generally

of the

After

.OZZkE

in regu-

and

Downstream

_--SNOCK

noz-

face

of Pb is

the expansion of the gases around shoulder C will continue unaffected

°"--_B-_]_

of their

role

pressure

__! FREESTREAMSURFACE

to these

annular

value

_,,/f_-_'_ /_I_

THROAT----x \

at

for an E-D

at the

The

ambient

ratios,

losses

an important

flow.

area

subject

WALL

__/I

sepa-

explained

to other Pb

plays

flow

because

not

in figure

nozzle

than

nozzles,

pressure

of the

large

ex-

(sea-

As

in thrust

applicable

centerbody

lating

with

Annular

special

before

_NOZZLE

MACHLINE_

may

ambient

occurs.

results

CONSTANT_

nozzles. gases

the

wall

I, for nozzles

this

the

operation)

nozzle

nozzle.

for annular

for bell

nozzle,

to pressures from

to those

or bell

pand ration

methods

and and

field

spike

secand

nozzle,

_Source: AIAA Paper No. 66-828, "Liquid Rocket Engines: Their Status and Their Future" By S. F. Iacobellis.

94

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

base achieved the

pressure through

and the base the secondary

pressure increase flow addition

requires a lengthy, detailed discussion; only the basic operation can be presented here. The primary flow (high-pressure gases) which produces the major portion of the engine thrust is exhausted from an annular-type combustion chamber and expands against the metal surface of the center tru-_ated-spike nozzle (fig. 4-18). The characteriF of the primary flow field upstream of the basv, shown as region 1 in figure 4-18, are determined by the annular throat geometry, the nozzle wall contour, and the ambient pressure. The annular primary flow continues to expand beyond the nozzle surface and encloses a subsonic, recirculating flow field in the base region (region 2). The pressure acting on the nozzle base contributes additional thrust to the nozzle. When a small amount of secondary flow is introduced into the base (added to the recirculating flow), the base pressure is increased further. As the secondary flow is increased, the overall nozzle efficiency (considering the additional flow) increases because of this increase in base pressure. There is a limit to this gain in efficiency, and an optimum secondary flow exists for each configuration. The outer surface of the annular primary flow is a free-jet boundary, which is influenced by ambient pressure. This ambient pressure influence on the primary nozzle flow endows this type of nozzle with altitude compensation. In operation at high-pressure ratios (i.e., altitude conditions), the outer free-jet boundary of the primary flow expands outward, governed by the Prandtl-Meyer turning angle at the throat. At low-pressure ratios (i.e., sea level operation), the relatively higher ambient pressure cornTOR_DJU._Ew

Figure

4-18.-Aerodynamic spike flow [ield i11usSrated under altitude conditions.

IDEAL

NOZZLE

(NO

LOSSE!

HIGH-AREA-RATIO AEROSPIK E NOZZLE

_

i, LL

/" /

L)

/ /

/_

i /

HICa-I-AR

/_

f

BELL

EA-RATIO

NOZZLE

I

I

,'

LU

/

//SEA'LEVEL

/

_

(VACUUM)

/

OPERATING

RANI

/ SEA

LEVEL

_,t_--

IJ

iI

50

100

I 200

IVACUUMI OPERATING

RANGE

] K JlJli{ 300

500

1000

2000

4000

PRESSURE RATIO (pc,/pa) Figure

4-19.-Nozzle

presses the flow field. pressure on the negative sure on the

perlormance

comparison.

outer free-jet boundary of the primary This compression increases the static the nozzle wall and partially offsets effect of the higher ambient presback side of the nozzle. The base

pressure also is increased with the higher ambient, because the compressed primary flow field, which influences the base pressure, has higher static pressures. This combination of flow field effects provides the altitude compensation inherent to the aerodynamic spike nozzle. Figure 4-19 presents the performance comparison of various nozzle designs. The nozzle thrust coefficient Cf for an ideal nozzle (i.e., a variable-area-ratio nozzle having the optimum expansion for each chamber pressure to ambient pressure ratio, pc/Pa) is shown together with those of the high-area-ratio aerodynamic spike and bell nozzle. As is evident, the CI curve of the aerodynamic spike follows the ideal nozzle performance (altitude-compensation), rather than dropping off rapidly like the bell design at low pc/Pa (i.e., sea level) operating points. M1 nozzles have a higher CI at a high Pc/Pa (i.e., vacuum). The development of the annular-nozzle concept may influence the design of rocket vehicles,

DESIGN OF THRUST CHAMBERSAND OTHER COMBUSTIONDEVICES

95

especially in the areas of boattail structure and mission staging optimization. The advantages and disadvantages of annular nozzles are summarized as follows:

4-23. The reader is urged to conduct his own calculations using the first stage as a guide, and to compare his results with those shown.

Advantages (1) Shortened nozzle length for the same performance, or increased performance (higher expansion area ratios) for a given length. (2) Improved performance at sea level or low altitudes. (Annular nozzles with high expansion area ratios can be used for a single-stage sea level to vacuum vehicle mission.) (3) The relatively stagnant region in the center of the nozzle can possibly be used for installation of gas generators, turbopumps, tanks, auxiliary equipment, and turbine gas discharges. (4) A segmented combustion chamber design approach can be used, easing development effort (individual segments can be built and tested during the early phases) and improving combustion stability.

Solution

Disadvantages (1) Relatively because surface

A-1 Stage Engine: From sample calculation Design

sea level

Substitute

Calculation

in some

(4-2)

Lay out the thrust chamber internal configuration (cylindrical combustion chamber with bell nozzle) for the engines on the Alpha vehicle with the data derived from sample calculation (4-1) and the following required chamber thrusts Ftc: engine:

Ftc

=747000

747 000 At - 1.531 x 1000 = 487 in:

Throat

diameter:

Dt = _=

24.9 in

Rt=_= Exit diameter:

De

--'-_V/_ -x

Use a combustion LO2/RP-1 (4-4):

24.9 =93.4

engine:

Ft%= 149500

lb at altitude

(c) A-3 stage

engine:

Ftc3=

Ib at altitude

(d) A-4 stage

engine:

Ftc4=

calculations

16000

chamber

application.

volume:

7500 lb at altitude and their results

are presented in the following for the first-stage engine only. For the other stages, the calculation results are summarized in figures 4-21 to

e= 14

(1-33):

area:

12.45

in

in

93.4 e =---_- = 46.7 in

L* of 45 in for

Substitute

into equation

Vc=487x45=21915cuin

Use a nozzle convergent half angle of 20 °, a contraction area ratio ec = 1.6, and a circular arc of radius R= 1.5Rt, or 18.68 in, for nozzle contour upstream of the throat. Chamber

diameter:

Dc = _'.'.6x

lb at sea level

(b) A-2 stage

The detailed

into equation

= 1000 psia;

Throat

Chamber

(a) A-1 stage

C/= 1.531; (Pc)as

R high cooling requirements, of higher heat fluxes and greater areas to be cooled.

(2) Heavier structural construction applications. (3) Manufacturing difficulties.

Sample

(4-1):

24.9 =31.5 R C

Use equation (4-7) to calculate convergent cone length Convergent

in

31,5 =_----=--=15.75

in

the chamber

cone length _ 12.45 (x/1.6 - 1) + 18.68 (sec tan 20 °

20 °- 1)

4.515 = 0.---3"64= 12.4 in

96

DESIGN OF LIQUID

PROPELLANT

ROCKET ENGINES

INJECTOR FACE

THRUST CHAMBER

AXrS

L n = 102.4"

Figure

4-20.-A-I

stage engine thrust (--14, 80% bell,

chamber L*=45",

internal (o=1.6.

-"-

configuration

INJECTOR THROAT FACE ,d

layout:

EXIT =9.25°

I

b I

8n=33 ° N

THRUST

J

__

_L_

14"R

CHAMBER AXIS

I

N0=5.94.

De=71"

L

I

17.3" - ! _ Figure

4-21.-A-2

stage

Ln =83.6"

engine _=40,

thrust

75% hell,

chamber,

internal

L*=26",

_c=1.6.

configuration

layout:

4mmmal

k.,

DESIGN

OF

THRUST

CHAMBERS

AND

OTHER

COMBUSTION

97

DEVICES

INJECTOR FACE

THRUST CHAMBER

AXIS

Figure

4-22.-A-3

stage

engine _=35,

thrust 70%

INdECTOR FACE

chamber,

bell,

L*=28",

internal

configuration

THROAT

EXIT

I

_1

V

layout:

_c=2.

Et:46.1,

`

THROAT CHAMBER

AXIS

Figure

4-23.-A-4

stage

engine _=35,

thrust 70%

bell,

chamber, L*=32",

internal ec=2.

configuration

layout:

98

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

Since

Using the frustrum cone volume equation and neglecting the slight rounding of the throat, the approximate convergent cone volume is obtained:

Volume =3 x12.4 [(15.75)2 + (12.45)2 +15.75×12.45

]

the calculations

for the thrust

chamber

configuration are based on the calculated design C[ value which has to be verified by later actual testing, a slight change of chamber pressure is usually allowed to compensate for C! deviations in order to meet the required thrust value.

= 7760 cu in 4.4 Required

volume

for cylindrical =21 915-

chamber section 7760:14 155 cu in

Required

length

for cylindrical chamber section = 14 155/1.6At= 18.17 inches

Distance

from injector face to throat = 18.17+ 12.40= 30.57, say 31 inches

Design an "80-percent bell" nozzle configuration using the parabolic approximation procedure. The nozzle contour downstream of the throat will be a circular arc of radius 0.382 Rt, or 4.75 inches. By definition, the nozzle length Ln will be 80 percent of the length for an equivalent 15 ° half-angle conical nozzle. Substitute into equation (4-7)

Ln=0.8×

[12.45

(VIT-1)+4.75 tan 15 ° (sec

15 °- 1!]

= 0.8 x 128 = 102.4

inches

The parabolic contour wall angles 0, and 0e can be derived from figure 4-14, for e= 14 and L/=0.8; On = 27.4 ° and 0e = 9.8 °. The location of N and E along the nozzle contour, with respect to throat and nozzle axis, can be calculated Nt =0.382

R: sin On = 2.19 inches

Na = Rt + 0.382 Rt(1 - cos 0n) : 12.99 inches Et : Ln = 102.4 Ea= Re=46.7

inches inches

With the aid of the established coordinates for points N and E, and the angles 0_ and 0e, a parabola can be fitted to complete the contour. The general layout of the A-1 stage engine thrust chamber is shown in figure 4-20. With the aid of a computer program, more accurate calculations of the divergent nozzle contour can be made by the method of characteristics.

THRUST

Techniques

CHAMBER and Their

COOLING

Selection

Because of the high combustion temperatures (4000 ° to 6000 ° F) and the high heat transfer rates from the hot gases to the chamber wall (0.5 to 50 Btu/in2-sec), thrust chamber cooling becomes a major design consideration. For shortduration operation (up to a few seconds), uncooled chamber walls can be used. In this case, the heat can be absorbed by the sufficiently heavy chamber wall material which acts as a heat sink, before the wall temperature rises to the failure level. For most longer durationapplications, a steady-state chamber cooling system has to be employed. One or a combination of the following chamber cooling techniques is often used: 1. Regenerative coollng.-Regenerative cooling is the most widely applied method and utilizes one or possibly both of the propellants, fed througt_ passages in the thrust chamber wall for cooling, before they are injected into the combustion chamber. (See par. 4.1 and fig. 4-1.) 2. Dump cooJing.-With this principle, a small percentage of the propellant, such as the hydrogen in a LO2/LH2 engine, is fed through passages in the thrust chamber wall for cooling and subsequently dumped overboard through openings at the rear end of the nozzle skirt. Because of inherent problems, this method has only limited application. 3. Film cooling.-Here, exposed chamber wall surfaces are protected from excessive heat with a thin film of coolant or propellant which is introduced through manifolded orifices in the chamber wall near the injector, and usually in several more planes toward the throat. The method has been widely used, particularly for high heat fluxes, either alone or in combination with regenerative cooling. 4. Transpiration cooling.-Transpiration cooling is accomplished by introducing a coolant

DESIGN OF THRUST CHAMBERSAND OTHER COMBUSTIONDEVICES

(either gaseous or liquid propellants) through porous chamber walls at a rate sufficient to maintain the desired combustion gas side chamber wall temperature. This method is essentially a special type of film cooling and has been widely used. 5. Ablative cooling.-In this process a sacrifice of combustion-chamber gas-side wall material is made by melting and subsequently vaporizing it to dissipate heat. As a result, relatively cool gases flow over the wail surface, thus creating a cooler boundary layer, assisting the cooling process. Ablative cooling has been used in numerous designs, initially mainly for solid propellant systems, but later equally successfully for low Pc, pressure-fed liquid systems. 6. Radiation cooling.-_tith this method, beat is radiated away from the surface of the outer thrust chamber wall. It has been successfully applied to low heat flux regions, such as nozzle extensions. The selection of the best cooling method for a given thrust chamber depends on many design considerations. There are no simple-and-fast rules. However, the following are the main factors which influence the selected design approaches: 1. Propellants.-The properties of the combustion products, such as temperature, specific heat, specific weight, viscosity, etc., have a direct bearing on the heat transfer rate and in turn affect the chamber cooling requirements and methods. The cooling properties of the propellants and their relative flow rate decide whether they are suitable or sufficient for regenerative or film cooling. Therefore, in evaluating a chamber cooling system, the propellants involved will be one of the primary design considerations. 2. Chamber pressure.-High chamber pressure is linked with higher combustion gas mass flow rates per unit area of chamber cross section and thus raises the heat transfer rate. Combined regenerative and film-cooling methods are usually employed for the stringent requirement of higher chamber pressure applications. 3. Propellant feed system.-The type of propellant feed used in an engine system determines the pressure budget for the system. In a turbopump-fed engine system, more pressure drop is usually available for chamber cooling. The

99

availability of this pressure drop permits the use of regenerative cooling which requires propellant pressure sufficient to force the coolant through the cooling passage before entering the injector. A pressurized-gas-fed engine system usually has more stringent pressure limitations and operates on relatively low chamber pressures. This suggests the application of film, ablative, or radiation cooling. 4. Thrust chamber configuration.-The geometric shape of the chamber affects local combustion gas mass flow rates and wall surface areas to be cooled. This influences the choice of cooling method. It can also arrangements for regeneratively wall thrust chambers. 5. Thrust

chamber

limit the design cooled tubular

construction

material.-The

properties of the thrust chamber materials will affect the cooling system design profoundly. Strength at elevated temperature, combined with heat conductivity properties of a metal, will Jetermine suitability for regenerative cooling systems. For film-cooled chambers higher allowable material working temperatures are desired to reduce heat transfer rates and thus film coolant flow rates. The application of radiationcooling to a chamber depends largely on the availability of high temperature (3000 ° F and up) refractory alloys. The success of ablative cooling depends entirely on the availability of suitable composite plastic materials. In practice, the design of thrust chamber cooling systems is a major link in the complete engine system design. It cannot be treated independently, without due consideration of other engine system aspects. For instance, optimization of the chamber pressure value for a highperformance engine system is largely limited by the capacity and efficiency of the chamber cooling system. In turn, chamber pressure will affect other design parameters such as nozzle expansion area ratio, propellant feed pressure, and weight. Because of the complex interrelation between these factors, the complete analysis of chamber cooling systems is a specialized field and requires thorough knowledge of heat transfer, fluid mechanics, thermodynamics, and thermal stresses. The engine system designer, therefore, will enlist the services of heat transfer specialists.

100

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

Gas-Side

Heat Transfer

One of the primary steps in the design of a thrustchamber cooling system is the analysisof the heat transferfrom the combustion gases to the chamber walls (gas-sideheat transfer).Because of the very high surface velocityof the gases along the chamber walls, the heat transfer occurs mainly through forcedconvection; i.e., throughthe transferof heat energy resultingfrom the relativemotion of different parts of a fluid. Before the gases can transferheat to the wall, the heat energy must pass througha layer of stagnantgas along the wall, calledthe boundary layer. The basic correlationforthiscomplicated convective heat transfercan be expressed by the followingequation: q : hg (Taw - Twg) where q = Heat flux or heat transferred across stagnant gas film per unit surface per unit time, Btu/in2-sec hg

=Gas-side heat transfer in2-sec-deg F

coefficient,

(4-10)

the area

dominantly influenced by the mass wlocity or the mass flow rate per unit area of the gas, subject to the exponent 0.8. In comparison, all other factors are relatively minor. A rough approximation of hg can thus be expressed by the following equation: hg=(p'V)

Twg = Hot-gas-sidelocal chamber-wall temperature,deg R The determination of the gas-side heat transfer coefficient hg is a rather complex problem. The convection phenomenon as it occurs in rocket thrust chambers eludes complete understanding. Attempts to compare analytical results with experimental heat-transfer data obtained on rocket thrust chambers have often shown disagreement. The differences are largely attributed to the initial assumptions for analytical calculations. For example, there is good evidence that oxidizing and reducing atmospheres covering a wide range of temperature exist locally in the combustion product gases within the thrust chamber, because of the imperfect mixing of the propellants at the injector face. This results in deviations from calculations based on the assumption of homogeneous product gases. However, it has been established by experiment _hat the beat-transfer coefficien_ is pre-

(4-11)

where p' = Free stream value of local gas density, lb/cu in = Free stream value of local gas velocity, in/set Thus, under normal circumstances, the heattransfer coefficient varies with the chamber pressure to the 0.8 power and throughout a given chamber inversely with the local chamber diameter to an exponent of 1.8. Based on experience with turbulent boundary layers, some relatively simple correlations for the calculation of the gas-side heat-transfer coefficient have been developed. A much-used form is that credited to Colburn

Btu/

Taw = Adiabatic wall temperature of the gas, deg R= (Tc)ns x turbulent boundary layer recovery factor (ranging from 0.90 to o.9s)

°8

Nu = C Re °8 Pr °_

(4-12)

where Nu = Nusselt number = hg D/k C =Dimensionless constant Re = Reynolds number = p'VD/_ = Free stream velocity,in/set Pr = Prandtlnumber = _Cp/k D =Hydraulic diameter,in k =Gas thermalconductivity,Btu/sec-in2deg F/in _t =Viscosity, Ib/in sec Cp=Specific heat at constantpressure,Btu/Ibdeg F or as Bartz

has shown

'++L+. + ,.'+c+++" It+ ] i+)o. (-+)°'o +,++ where R = l_dius of curvature throat, in

of nozzle

o = Correction factor for property across the boundary layer

contour variations

at

101

DESIGN OF THRUST CHAMBERSAND OTHER COMBUSTIONDEVICES

A = Area under consideration along chamber axis The value of a can be evaluated in terms of nozzle stagnation temperature, local gas-side chamber wall temperature, and local Mach number. 1

values of a for various Twg/(Tc)n s and y, as computed by Bartz, are shown in figure 4-24. If Pr and # data are not available for particular combustion gas mixtures, the following equations can be used for approximate results:

(4-15)

Pr=9__5 /_=(46.6 where T= temperature Equations (4-13),

× 10-_°) _ °ST °'6

(4-16)

of gas mixture, °R (4-14), (4-15), and (4-16)

can be used to calculate the ues along the thrust chamber the calculated values can be lower than the actual ones if ditions exist:

approximate hg valwalls. However, expected to be the following con-

0.6

"'l O -I'z

I J

o.6[ 1 4

3

TI/

I

JI

I

I

I

I

_1

2

I

Z

3

4

Figure 4-24.-Values property variation

The calculated values may be higher than the actual ones, because of the following: (1) The combustion reactions may not be completed in the chamber. (2) The combustion gases may deposit solids on the chamber walls, which act as insulators. In certain propellant combinations, the combustion products contain small amounts of solid particles. These solids tend to deposit on the chamber wails, and form a rather effective insulating layer. A quantitative evaluation of _he insulation effectiveness of this layer, necessary for correct heat transfer calculations, has been accomplished only experimentally. In the case of the LO2/RP-1 combination, carbon solids are deposited on the chamber walls. After a firing, the carbon gives the interior of the thrust chamber the appearance of being freshly painted black. The outer surface of the carbon appears sooty and can easily be removed by light rubbing. Underneath the exterior soot layer is a harder, graphitelike layer which can also be removed, but is more tenacious. This carbon deposit significantly increases the gas-side thermal resistance. The temperature of the carbon deposit at the hot gas-side interface approaches the gas temperature as the carbon thickness increases. The values of the thermal resistance of the carbon deposit based on actual experimental testing results of a thrust chamber burning LO2/RP-1 are shown in figure 4-25. For the heat transfer calculation of the gasside heat transfer with solid deposit on chamber walls, the following equations can be used q = hgc (Taw - Twg)

I

CONTRACTION

(1) A substantial fraction of the combustion gases are strong radiators. (2) There is substantial dissociation, with subsequent recombination near the wall. (3) There are strong high-frequency flow instabilities.

5 6

I 78510

20

30

40

where hgc = overall gas-side Btu/in2-sec-deg F

I[XP&NSION

of correction factor a for across boundary layer.

thermal

1 hgc-

conductance,

(4-18)

i hg +

(4-17)

Rd

102

DESIGN

!,!

2400

OF

LIQUID

PROPELLANT

,;

ROCKET

ENGINES

= 6460 × (0.975) 2 : 6140 ° R (See eq. 1-32a and 1-41). From sample calculation

(4-1):

22(311 _

Design I_ u? ,8o0

From sample

c* = 5660 ft/sec

calculation

(4-2):

i

Dc=24.9 Mean radius fzoo

of the throat

I

2 CONTRI.CllON _

I

2

4 t _

_¢ AREA

I ,

6 8 EXPANSION

_

,

i

12

- 11.71 in

e 44)

yR 1.222 x Cp - (),_ I),I - (i. 222 - I) × 778 = 0.485 Btu/lb-deg

From equation where Rd thermal resistance caused by the solid deposit, in2-sec-deg F/Btu When there is no solid deposit, Rd =0 and hgc = hg, and equation (4-10) is used for heat transfer calculations.

Calculation

contour 18.68+4.75 2

_4

RATIO

Figure 4-25.-Thermal resistance el carbon deposit on chamber walls LO2/RP-1, mixture ratio = 2.35, (Pc)n s = 1000 psia.

Sample

in

(4-15):

4 × 1.222 Pr : (9 × 1.222) - 5 =0.816

From equation

(4-16):

: (46.6 × 10 -1°) × (22.5) 05 x (6140) 0.6

(4-3)

Determine the approximate design gas-side overall thermal conductance hgc in the combustion chamber, at the throat, and at the exit nozzle point of e=5, for the regeneratively cooled thrust chambers on the A-1 and A-2 stage engines.

= 46.6 × 10-1° × 4.76 × 188 - 4.18 x 10 -6 lb/in-sec From equation

h -V0"026

g-L

(4-13):

x((4"18×10-_)°2x0'485) 0.8160.6

Solution

0.9

(_.) A-I Stage Engine First, let us consider equation (4-13). The combustion reactions are assumed to be homogeneous and complete. From figure 4-3 the following values are derived for the chamber product gases, for LO2/RP-1 mixture ratio of 2.35: (Tc)ns

F

at (Pc)ns = 1000 psia

and a

lb/mol,

y= 1.222

(Tc)ns

=Theoretical

(Tz)ns

: 0.01366

× (c* correction

factor) 2

7

At\ 0.9 × 0.046 × 4.02 × 1.078 × k-_} o

"At_O.9

= 0.0027 x ("_')

= 6000° F or 6460 ° R, =22.5

The design

×\-g-6-65

a

Since the carbon deposit temperature approaches the gas temperature, a (Twg/(Tc)ns) value of 0.8 is used to determine the a values from figure 4-24 (),_1.2). At the combustion chamber:

J

(7

DESIGN OF THRUST CHAMBERS AND OTHER

COMBUSTION DEVICES

b(b_)A-2 (__)

°9:(1_16) " 1" 0.9 :0,655,

a=l,05

Again,

Stage the

Engine

combustion

to be homegeneous hg : 0.0027

4-4,

× 1.05 = 0.00185

x 0.655

Btu/in2-sec-deg

F

the

= 800 At the

following

chamber

product

psia

and

103

reactions

and

complete.

values

are

gases,

assumed figure

derived

for

for LO;/LH2

a mixture

ratio

the

at (Pc)as

of 5.22:

throat: (Tc)ns

= 5580 ° F or 6040

° R, =12

-_)°'9=l,a=

hg = 0.0027

x 1 x 1 : 0.0027

Btu/in

2-sec-deg

design

exit

nozzle

point

(Tc)ns

=

(Tc)ns

x (c*

From

sample

calculation

× 0.8 : 0.000507

The experimental used to determine Rd,

for the

resistances

Btu/in2-sec-deg

data of figure 4-25 the values of thermal carbon

factor)

× (0.975)

2

2 = 5740 ° R

(4-1):

a=0.8 Design

× 0.235

correction = 6040

of

e=5,(_-)°'9=(1)°9:O.235,

hg = 0.0027

y= 1.213

F Theoretical

At the

lb/mol,

1 The

ance

are From

deposit.

The

F

From

figure

c* = 74S0

ft/sec

4-21:

Dt = 11.2 in

can be resistMean

thermal

radius of the throat contour =

are 8.4+2.14 - 5.27 in

in2-sec-deg 1670

F

in:-sec-deg

Btu

,

1125

2

F

Btu

' in2-sec-deg

and

for points and

the

at the exit

Substitute combustion

combustion

nozzle into

area the

chamber, ratio

equation

Cp = (y_

Btu

the

yR

From

throat,

1.213

equation

(4-18);

Btu/in2-sec-deg

F

1.21a

equation

- 5 = 0.820

(4-16):

F

1670

= (46.6

x 10-lo)

(12) o.s (5740)o6

x 10- ,o × 3.47

x 180

throat =2.92x

hg c -

At the

1 --+ 0.0027 exit

= 0.00067

Btu/in2-sec-deg

nozzle

1 -0.000507

From

lb/in-sec

equation

(4-13):

of e= 5.

= 0.000276 e1645

10 -6

F

1125

hg=L [ 0.026 hg c-

Btu/lb-deg

(4-15): 4x

at the

From = 0.00045

0.943

of e = 5.

= 46.6 at the

- 1) × 778-

Pr = (9 x 1.213)

1 --+ 0.00185

1544 i----_

x

1)g = (1.213

chamber

1 hgc -

1645

F

Btu/in2-sec-deg

x ((2.92x,

_10-6)°2×0.943)

.

F x\

7480

x\5--._

]

j\-_--]

a

104

DESIGN OF LIQUID PROPELLANT

(A_

ROCKET ENGINES

t.,4_CHAMBER

09

:0.01605 ×0.0828 ×2.69×1.078×\--X-J ° GAS SIDE BOUNDARY

I INNER WALL ['_qI.--COOLANT SIDE BOUNDARY LAYER

LAYER

I A_ x°9 _At = 0.00385

Since wails,

x _---_-)

there

is

no solid

an average

and

(1500/5740)

or 0.26

values

figure

At the

--

0.9

on the

wall

temperature

of

a (Twg/(Tc)ns)

is used

the

a

Tw_

chamber:

G.__16

)

0.9

= 0.655,

RADIAL

a= 1.38

At the

x 0.655 x 1.38 = 0.00348 Btu/in2-sec-deg

Figure

--Twc

DISTANCE

CENTER hgc = hg : 0.00385

COOLANT

°1 It

valueof

to determine

\

To. COM__

chamber

4-24.

combustion

(__)

deposit

gas-side

1500 ° R is assumed, from

Z

a

OF

4-26.-Heat

FROM

CHAMBER

trans[er

F

erative

Tco

schematic

for regen-

cooling.

throat: =hc(Twc-Teo)

_)°9

fig c :hg=0.00385

Btu/in2-sec-deg

(4-21)

1

H-

× 1 x 1.35 = 0.00520

-Tco)

= H (Taw

= 1, o=1.35

(4-20)

1

t

(4-22)

1

F where

At the

exit

nozzle

point

of e = 5:

(-_)°9=(1)°9=0.235,

q

= Heat

hgc

= Overall

Btu/in2-sec

gas-side

thermal

Btu/in2-sec-deg

a=1.16 hc

hgc = hg = 0.00385

flux,

× 0.235 × 1.16 =0.00105 Btu/in2-sec-deg

out deposits, =Coolant side

k

= Thermal

t

Cooling

conductivity

= Chamber

wall

Taw =Adiabatic The

heat

chamber tween

transfer

can two

partition. matically.

be described fluids,

Figure

4-26

The transfer

through

the

by the

as

moving

of heat

chamber

in a regeneratively

layers,

walls, following

shows

general from

this

heat

which

process

include

equations:

can

be-

sche-

correlation

combustion

coolant

flow

a multilayer

steady-state the

to the

the

through

cooled

deg

metal

be expressed

with-

coefficient,

of chamber

wall,

F/in thickness,

in

temperature

of the

gas,

R

Twg= Gas-side Two = Coolant

wall temperature, deg R side wall temperature, deg

Teo H

bulk temperature, deg heat-transfer coefficient,

= Coolant = Overall

gases the

wall

4-18;

F

Btu/in2-sec-deg Regenerative

eq.

hg c = hg) heat-transfer

Btu/in2-sec-deg F

conductance,

F (see

sec-deg The

bulk from

cooling

passages,

absorbed,

the

and

balance

of these

chamber

walls

Btu/in:-

F

temperature

creases

R

R

Too

of the

point

of entry

as

a function

of the

coolant

parameters, at temperatures

coolant

until

of the

flow

the

heat

rate.

Proper

to maintain below

in-

it leaves

those

the at

DESIGN

OF THRUST

CHAMBERS

which failure might occur because of melting or stress, is one of the major criteria for the design of regeneratively cooled thrust chambers. For metals commonly used in thrust-chamber walls, such as stainless steel, nickel, and Inconel, the limiting hot-gas-side wall temperature is around 1500°-1800 ° F. The resultant differences between combustion gas temperature and wall temperature range from 2500 ° to 6000 ° F. Assume a station in the thrust chamber with gas temperature Taw and coolant bulk temperature Tco. Referring to equation 4-21, it is seen that the heat flux q, which must be the same through all layers, is a function of the temperatures, and of overall heat transfer coefficient H. The value of H is composed of the individual coefficients for the boundary layers and the chamber metal wall (eq. 4-22). The smaller H, the smaller is q. However, it is one of the major design goals heat transfer

to keep coefficient hgc low, but coefficient hc and conductivity t/k

high, in relation to hgc. Since the temperature differentials are inversely proportional to the heat-transfer coefficients of the heat flow paths, the temperature drop will then be steepest between hot gas and inner chamber wall. The effect is analogous to voltage drops along resistors in electrical circuits. It is noted that the heat absorbed by the propellant used for regenerative cooling raises temperature of the propellant, and thus the energy level before it is injected into the combustion chamber. However, this effect on overall engine performance is slight, the gain usually being less than 1 percent. On the other hand, regenerative cooling with attendant pressure losses requiring additional turbopump power or higher gas pressurization levels imposes a performance penalty.

AND

OTHER

COMBUSTION

105

DEVICES

The characteristics

of coolant

side

heat

transfer depend largely on the coolant pressure and coolant side wall temperature. In figure 4-27, the heat flux is plotted versus wall temperature for a constant coolant pressure, bulk temperature, and flow velocity. Curve A indicates the behavior of heat transfer at coolant presst, res below critical. Line segment At-A2 represents the heat transfer without boiling when the wall temperature is below the saturation temperature of the coolant corresponding to the fluid pressure. As the wall temperature at A: exceeds the saturation temperature by a certain margin (50 ° to 100 ° F), bubbles will form within the coolant layer close to the wall. The bubbles grow continuously out into the colder liquid stream until condensation at the vapor to liquid surface begins to exceed the rate of vaporization at the base of the vapor bubble, whereupon the bubbles start to collapse. This process, which occurs at high frequencies, is described as "nucleate boiling." It substantially increases the heat-transfer coefficient, resulting in little increase in wall temperature for a wide range of heat fluxes. The heat transfer with nucleate boiling is represented by line A2-A 3. At A 3, further increase in the heat flux abruptly leads to such'a dense bubble population that the bubbles combine into a vapor film with an attendant large decrease in heat-transfer coefficient. The region of heat transfer with film boiling is represented by line A3-A 4. The resulting increase in

_6

m

.A 3

-w,.

....

° .........

¢o P"_

x

_

C U RVE

A

(Pco =I/2

N

Pc_,r,cAO

z4

Coolant Side Heat Transfer The coolant side heat-transfer coefficient hc is influenced by many factors. At the high heat fluxes and temperatures encountered in thrust chamber operation, the propellants used for cooling may become corrosive, may decompose or deposit impurities upon the heated surface, thereby reducing cooling effectiveness. It is impossible to calculate the hc values under these conditions without experimental data.

Mr At _

_"

O B_2

I

I I

4

6

,"--CURVE ( P"OI

e

I0

> I PCmTICALI

12

14

I)

16

is

20

COOLANT SIDE WALL TEMP. TWC (°F) X I0 -z Figure 4-27.-Heat flux versus coolant side wall temperature of typical propellant in various heat trans/er regions.

106

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

wall temperature is so high that failure of the wall material often occurs. The heat flux at A_ is defined as the upper limit of nucleate boiling of the coolant qul, which therefore should be used as the design limit for a regenerative cooling system. Curve B indicates the heat transfer behavior of a coolant above critical pressure. Since no boiling can occur, the wall temperature continually increases with increasing heat flux. Line B1-B2 represents the heat-transfer region, when the wall temperature is below the coolant critical temperature. The heat-transfer coefficient remains essentially constant. As the wall temperature reaches the critical temperature B2 and higher, a gradual transition to a stable supercritical vapor-film boundary layer begins, which results in somewhat lower heat-transfer coefficients. Line B2-B 3 represents the heat transfer in this region. Wall failure temperatures are usually reached at lower temperatures when the coolant is above the critical presstire than when it is below it. Where possible, a coolant operating pressure between 0.3 to 0.7 of critical pressure should be used to take advantage of the high heat-transfer coefficients available with nucleate boiling. However, in most systems, particularly those which are fed from a turbopump, the cooling jacket pressure, which is equal to or larger than the sum of chamber pressure and injection pressure, is supercritical. For the nonboiling subcritical temperature regions of both, subcritical and supercritical coolant pressures (AI-A2 and B_-B 2 in fig. 4-27), the relationship between wall temperature and heat flux, which depends on the heat transfer coefficient hc, can be predicted with sufficient accuracy for design purposes with the help of the Sieder-Tate equation (eq. 4-23) for turbulent heat transfer to liquids flowing in channels:

gw =coolant viscosity at coolant sidewall temperature d = coolant passage hydraulic diameter, in k = coolant thermal conductivity, Btu/sec-in deg F/in p =coolant density, ib/in 3 Vco = coolant velocity, in/sec Cp = coolant specific heat at constant pressure, Btu/ib-deg F The heat flux at the upper limit of nucleate boiling qul can be estimated from qul qnonboiling

Nu=C 1 Re °SPr°4

(4-23)

(ju+)

where C I =a constant (different values coolants) Nu =Nusselt number= hcd/k

for various

Re = Reynolds number = pVcod/_ Pr = Prandtl number = i_Cp/k /_ =coolant viscosity at bulk temperature

C 2 x 10* (4-24)

PcoG

where C2

= constant, coolant

qnonboiling

-- heat

Pco G

= coolant pressure, psia = coolant maximum flow rate unit area, lb/in2-sec

its value used

flux without Btu/in2-sec

depending nucleate

on

boiling,

per

When the heat is transferred through a vaporfilm boundary layer (coolant at supercritical pressure and temperature, region B2-B 3 in fig. 4-27), the coolant-side heat-transfer coefficient hc can be estimated from

hc: 0.029 Cpp °2 (.GO*,_.2._, _ oss pr2/3 \--d'_] \Twc!

(4-25)

where Cp

Pr G

= coolant specific heat at constant pressure, Btu/Ib-deg F coolant viscosity, Ib/in-sec = Prandtl number = coolant weight flow rate per unit area, Ib/in2-sec

coolant passage hydraulic diameter, in = coolant bulk temperature, deg R Twc--coolant side wall temperature, deg R The bulk temperature of most coolants should be kept below the critical temperature, since the vapor-film heat-transfer coefficient would be too d

014

2-

z

Tco

low to cool the wall effectively. The cooling capacity of the liquid-state regenerative coolant system can be estimated by Qc

= #cCp

(Tcc - Tci)

(4-26)

DESIGN OF THRUST CHAMBERSAND OTHER COMBUSTIONDEVICES

k._.._,

_f.._

transfer calculation. There are several basic design approaches for regenerative-cooled thrust chambers. Axial-flow cooling jackets, made up of tubes, are used in the design of large thrust chambers (3000 pounds of thrust and up); coaxial shells separated by helical ribs or wires are typical of the smaller thrust chamber designs. Figure 4-1 shows a large regenerative cooled tubular wall thrust chamber. Figure 4-28 represents a typical coaxial shell design for a smaller thrust chamber.

L

i

In this design, the coolant passage is defined as the rectangular area between inner and outer shell and two adjacent ribs, which are wrapped helically around the inner shell or liner.

i

/ Figure

4-28.-Coaxial

shell

way. Note overheated on chamber wall.

thrust and

Tubular

chamber

cuta-

burnt-through

spot

where Qc =coolant We = coolant

107

capacity, Btu/sec mass flow rate, lb/sec

Cp

=coolant specific heat at constant pressure, Btu/lb-deg F Tcc =coolant critical temperature, deg R Tci = coolant inlet temperature, deg R The allowed value of the total chamber wall gas heat-transfer rate Q should be kept below Qc by a safe margin (Q V2, due to friction losses. Ideally, the gas should leave the blades at very low absolute velocity C 2 and in a direction close to axial for optimum energy conversion in the blades. The forces generated at the rotor blades are a function of the change of momentum of the flowing gases. The following correlations may be established for design calculations of the rotor blades of a single-stage, single-rotor turbine.

(6-129)

For subsequent calculations, relation will be useful:

Axial

(6-128)

,TdmN 720

and

The velocity vector diagram shown in figure 6-56 describes graphically the flow conditions at the rotor blades of a single-stage, single-rotor turbine, based on tile mean diameter din. The gases enter the rotor blades with an absolute velocity C_, and at an angle a, with the plane of rotation. The tangential or peripheral speed of the rotor blades at the mean diameter is U. V 1

Tangential force gas flow/sec):

cos E1 +V2 cos/_2)

_b =

cos22a_(1

is some reaction

component

of C_.

' cos_ + _b co--_-_-_fll)(6-134)

or expansion

gas flowing through the blades, flow velocity at the rotor blade calculated as

of the

the relative gas outlet can be

of gas V 2=\/kb2Vl2 + 2gJ_?nAHi-2'

(6-135)

DESIGN OF TURBOPU_P PROPELLANT-FEED SYSTEMS

Amount of reheat gas flow:

qbr=(1

-kb

in the rotor

v?,

)2g j-r(1-

blades,

Btu/lb

r]n)A/-/l-2'

of

(6-136)

where al,

a2 =absolute gas flow angles at the inlet and outlet of the rotor blades, dog ill, fi2 = relative gas flow angles at the inlet and outlet of the rotor blades, dog C,, C: =absolute gas flow velocities at tile inlet and outlet of the rotor blades, ft/sec V,, U dm _?n

V 2 =relative

gas flow velocity

at the inlet

and outlet of the rotor blades, ft/sec = peripheral speed of the rotor, ft/sec = mean diameter of the rotor, in = equivalent nozzle efficiency applicable to the expansion process in the blades

AH,_2,=isentropic enthalpy drop of the gases flowing through the rotor blades due to expansion or reaction, Btu/lb; AH1-2' = 0 if only impulse is exchanged All parameters refer to the mean diameter d m, unless specified otherwise, The turbine overall efficiency Ut defined by equation (6-19) can be established for a single-stage, single-rotor impulse turbine as y_: _?n_b _rn

(6-137)

where rl_ =nozzle efficiency )7b =rotor blade efficiency fir. =machine efficiency indicating the mechanical, leakage, and disk-friction losses in the machine. Equation

(6-134)

shows

that

the blade

effi-

ciency ;?b improves when/32 becomes much smaller than I3L. Reduction of/32 without decreasing the flow area at the blade exit can be achieved through an unsymmetrical blade design (fig. 6-56), where the radial blade height increases toward the exit. In actual designs, the amount of decrease of fi2, or the increase of radial height, is limited considering incipient flow separation and centrifugal stresses. Generally, the _2 of an unsymmetrical blade will be approximately f3_-(5 ° to 15°). Equation (6-134) also indicates that _b improves as a_ is reduced.

1L

243

Design values of kb vary from 0.80 to 0.90. Design values of _b range from 0.7 to 0.92. Referring to figure 6-56, the radial height at the rotor inlet, hb, is usually slightly larger (5 to 10 percent) than the nozzle radial height hn. This height, together with the blade peripheral speed U, will determine the centrifugal stress in the blades. The mean diameter of the rotor blades is defined as dm= d_- hb, where d t is the rotor tip diameter. Pitch or blade spacing, Pb, is measured at the mean diameter dm. There is no critical relationship between blade pitch Pb and nozzle pitch Pn. There just should be a sufficient number of blades in the rotor to direct the gas flow. The number of blades zb to be employed is established by the blade aspect ratio, hb/Cb and the solidity Cb/Pb, where Cbis the chord length of the rotor blades. The magnitude of the blade aspect ratio ranges from 1.3 to 2.5. Design values of blade solidity vary from 1.4 to 2. Best results will be determined by experiment. The number of rotor blades should have no cdmmon factor with the number of nozzles or of stator blades. The blade face is concave, with radius rt. The back is convex, with a circular arc of small radius rr concentric with the face of the adjoining blade ahead. Two tangents to this arc to form the inlet and outlet blade angles 0b_ and Oh2 complete the blade back. The leading and trailing edges may have a small thickness tb. The inlet blade angle 0b_ should be slightly larger than the inlet relative flow angle fl_. If Ob_ fi_, the stream will strike the concave faces of the blades and tend to increase the impulse. The outlet blade angle Oh2 is generally made equal to the outlet relative flow angle /32. The mass flow rate _'t through the various nozzle and blade sections of a turbine is assumed constant. The required blade flow areas can be calculated by the following correlations. Note that the temperature values used in calculating the gas densities at various sections must be corrected for reheating effects from friction and turbulence.

Wt =

p1VlAblebl 144

_p2V2Ab2eb2 144

(6-138)

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

244

Total

blade

inlet area,

Abi =Zbbbzhbz

in2:

=Zbhbi (Pb sin Obl -tb)

(6-139)

Total blade exitarea,in2: Ab2=Zbbb2hb2=Zbhb2(Pb

sin Ob2-tb) (6-140)

where Pb

= pitch

P,, P2

=density outlet

V z, V 2

=relative gas flow velocities at the inlet and outlet of the rotor blades, ft/sec

ebl, (b2

=

area coefficients at inlet of the rotor blades --number of blades

and outlet

and outlet

Zb

or rotor blade spacing = rrdm/zb, in

(6-140a)

of the gases at the inlet and of the rotor blades, lb/ft _

hb_,

bb2

:radial height at the inlet of the rotor blades, in

bbl,

bb2

:

passage widths (normal to flow) at the inlet and outlet of the rotor blades,

in

0b_, 062 :rotor blade angles at inlet and outlet, deg tb = thickness of blade edge at inletand outlet,in Typical constructionsof rocketturbinerotor blades and disks are shown in figures6-53,6-55, 6-56,and 6-57. Usually,blades are designed with a shroud, to preventleakage over the blade tips and to reduce turbulenceand thus improve efficiency.Frequentlythe shroud forms an integralportionof the blade, the shroud sections fittingcloselytogetherwhen assembled. In otherdesigns the shroud may form a continuous ring (fig.6-55)which is attachedto the blades by means of tongues at the blade tip,by rivets, or is welded to the shrouds. The blades may be eitherwelded to the disk, or attached to itusing "fir-tree" or otherdovetailshapes. The main loads to which a rotorblade is

of various blade sections at different radii generally do not fall on a true radial line. Thus the centrifugal forces acting upon the offset centroids will produce bending stresses which also are a maximum at the root section. 2. Bending due to gas loading.-The tangential driving force and the axial thrust produced by the momentum change of the gases passing over the blades may be treated as acting at the midheight of the blade to determine the amount of bending induced. 3. Bending due to vibration loads.-The gas flow in the blade passages is not a uniform flow as assumed in theory, but varies cyclically from minimum to maximum. The resultant loads represent a dynamic force on the blades, having a corresponding cyclic variation. If the frequency of this force should become equal to the natural frequency of the blades, deflections may result which will induce bending stresses of considerable magnitude. Detail stress analyses for rotor blades can be rather complex. A basic approach is to counteract a major portion of the bending moments from gas loading with the bending moments induced by the centrifugal forces at nominal operating speeds. This can be accomplished by careful

i-

INTEGRAL

_

TYPE

SHROUD

gLADE _

ROOT A-A

S.HROUDED BLADES CASTING PROCESS

FABRICATED

BY PRECISION

exposed can be dividedintothreetypes: _ BLADE

1. Tension and bending due to centri[ugal [orces.-The radial component of the centrifugal forces acting on the blade body produces a centrifugal tensile stress which is a maximum at the root section. As a remedy, blades are often tapered, with the thinner section at the tip, for lower centrifugal root stresses. The centroids

JOINT

L.,_

!_

z

DISK

BLADES WELDED THE DISK

TO

8LADES TO THE

ATTACHED DiSK BY

TYPICAL "FIR TYPE TANG

TREE"

"FIR TREE"TYPE TANG

Figure

6-57.-Typical

rotor blade

constructions.

DESIGN OF TURBOPUMP

blade

design.

stresses

Thus

become

design,

while

location

and

later ing

the

other root

details

to fulfill

design are

where

of blade

such

as

centroid

The

tensile

stress

of uniform

cross

at the

most

duced

section,

follow-

blade

root

from the torque. As seen in figure 6-55, turbine disks are generally held quite thick at the axis, but taper off to a thinner disk rim to which

root

section

psi:

blades tions, both

Sc = 0.0004572

stresses in a turbine rotor disk are inby (1) the blades, and (2) the centrifugal

addition, there will be shear stresses resulting

critical.

at the

245

SYSTEMS

forces acting on the disk material itself. In

established

requirements. are

The

in blade

are

established

stresses

Centrifugal

tensile

consideration

configuration

correlations

section

centrifugal

a first

PROPELLANT-FEED

1pbhbdmN2 g

(6-141)

are

attached.

it is

possible

radial

at all

and

points,

rotor

In single-rotor to design

tangential

shear

applications,

cause

of the

it is

greatly

a disk

stresses

being

the

applicaso that are

neglected. difficult

uniform In multi-

to do this

increased

axial

be-

length

and

Centrifugal tensile stress at the root section of the resulting disks.

a tapered blade, psi:

large

Equation Sct=

O.O0045721pbhbdmN

2

(6-144) in a uniform

ing

blade

rotor

moment

section,

in-lb:

due

to gas

loading

at the

root

(6-143)

tensile turbine

=turbine

tr of the

blade

hb

blade

height,

=average

din =mean

diameter

=turbine

speed,

ar

= sectional

area

sectional

area

_i,t = turbine

gas

=number

of blades

Ft

= tangential

force

lb/lb/sec

(eq.

The culated

bending from

vibration design rate

data.

stresses forces

blade

root,

at the

blade

tip,

rate,

lb/in

s

in

rpm

of the disk at the axis, in

= thickness of the disk rim at d d, in (6-144a) permits estimation of the

in 2

Sd = O.O0044251W

on

the

blades,

Sd

= centrifugal

(6-127)) on

disk, the

blades,

lb/lb/

(6-144a)

at the be

blade root.

obtained caused blades.

root

can

bending

moment. from

is

fitted force The

with

The past

total

stress

by adding

these

by the

disk ad =disk N

centrifugal

from

the

of gravity

axis,

speed,

good

turbine

that

at maximum

at

the

S d calculated 0.75

of the

turbine

lb

center

to 0.8

Turbine high-temperature

rotor

of tile half

in

cross-sectional

addi-

about

stress

disk,

of the

=turbine For

a sepa-

produces

of the

=distance

be cal-

estimated

tensile psi

Wd = weight ri

centrifugal

on the

dri N2 ad

g

in 2

lb/sec

acting

can

is

to those

material,

where

resultant

at the

section acting

at the

acting

If the its

stresses root

in

stresses

the

of a constant

psi

disk,

speed,

Equation

3

(6-131))

stresses

shroud,

tional the

(eq.

(6-144)

stresses in any turbine disk, neglecting effects of the rotor blades:

rotor,

flow

thrust

sec

the

neglect-

rpm

zb

Fa = axial

lb/in

in

of the

N at'=

material,

stress

disk

of the

to =thickness

pb =density

disk,

stator

dd 2N2 log ___o_

disk,

of the

d d = diameter N

where

to estimate

turbine

and

where

Pd= density 2

be used

stress

Sd=O.OOOll41pd

stress _ hbcct /F,_+Fa S-2Zb _

rotor

effects:

Sd =centrifugal

M

between

may

stresses

(6-142)

Bending

gaps

area,

in 2

rpm design,

it is

allowable

design

by equation material

recommended rotating

(6-144a)

yield

blades

and

alloys

of three

disks

speed,

should

be

strength. are different

made

of

base

246

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

materials: iron, nickel, and cobalt, with chromium forming one of the major alloying elements. Tensile yield strength of 30 000 psi minimum at a working temperature of 1800°F is an important criterion for selection. Other required properties include low creep rate, oxidation and erosion resistance, and endurance under fluctuating loads. Haynes Stellite, Vascojet, and Inconel X are alloys frequently used. The rotor blades are fabricated either by precision casting or by precision forging methods. Rotor disks are best made of forgings for optimum strength.

Design of Single-Stage, Two-Rotor VelocityCompounded Impulse Turbines (figs. 6-9, 6-55, and 6-58) In most impulse turbines, the number of rotors is limited to two. It is assumed that in a singlestage, two-rotor, velocity-compounded impulse turbine, expansion of the gases is completed in the nozzle, and that no further pressure change occurs during gas flow through the moving blades. As mentioned earlier, the two-rotor, velocity-compounded arrangement is best suited for low-speed turbines. In this ease, the gases ejected from the first rotor blades still possess considerable kinetic energy. They are, therefore, redirected by a row of stationary blades into a second row of rotor blades, where additional work is extracted from the gases, which usually leave the second rotor blade row at a moderate velocity and in a direction close to the axial. The velocity diagrams of a single-stage, tworotor, velocity-compounded impulse turbine are shown in figure 6-58, based on the mean rotor diameter. The peripheral speed of the rotor blades at this diameter is represented by U. The gases leave the nozzles and enter the first rotor blades with an absolute velocity Cx, at an angle a_ with the plane of rotation. V_ and V2 are the relative flow velocities in ft/sec at the inlet and outlet of the first rotor blades, The gases leave the first rotor blades and enter the stationary blades at an absolute flow velocity C 2, and at an angle a2. After passing over the stationary blades, the gases depart and enter the second rotor blades at an absolute flow velocity C 3, and at an angle as. V3 and V4 are the relative inlet and outlet flow velocities at the second rotor

blades. Angles ]3,, f12, /33, and /34 represent the flow directions of V,, V 2, V 3, and V 4. As with single-rotor turbines, the exit velocity from any row of blades (rotary or stationary) is less than the inlet velocity, because of friction losses. It can be assumed that the blade velocity coefficient any row of blades:

k b has the same value

V2 _Ca _V4 kb- v 1 C: V 3

for

(6-145)

In a multirotor turbine, the total work transferred is the sum of that of the individual rotors:

C]

Vl

NOZZLE _lI

ROTA --

-_

....

FIRST

ROTOR

C2

ROT.T,ONOX "_'_

_/,,_,=

SECOND ROTOR

4

S_a4 _¥4 U

Figure 6-58.-Velocity diagrams o{ a typical single-stage, two-rotor, velocity-compounded impulse turbine.

Total

work transferred

rotor turbine, U E2b=-_(Cl

ft-lb/lb

to the blades

of a two-

of gas flow/see

cos al +C2 cos a2 +C

a

COS

a3+C

4

C0S

a4)

H

=g(V,

cos/31 +V2 cos/32 +V 3 cos /3a +V4 cos /34)

(6-146)

DESIGN OF TURBOPUMPPROPELLANT-FEED SYSTEMS

Combined nozzle and blade efficiencyof a tworotorturbine: E2b T]nb= JAH

(6-147)

where AH =overall isentropicenthalpydrop of the turbinegases, Btu/Ib = totalavailableenergy contentof the turbine gases (eq.6-17) Equation (6-137)can be rewrittenforthe turbine overallefficiencyqt of a two-rotorturbine as

7It = T/nbT/m

(6-148)

Ideally,tin b is a maximum forthe singlestage,two-rotor,velocity-compoundedimpulse turbinevelocityratio

U

C,

cos

a

I

4

i.e.,when U= ¼C,t. The workload forthe second rotorof a two-rotor, velocity-compounded turbineis designed at about one-fourthof the totalwork.

The design procedures for the gas flow passages of the rotor and stationary blades of a single-stage, two-rotor turbine are exactly the same as those for a single-rotor turbine. However, velocities and angles of flow change with each row of blades. As a result, the radial height of symmetrical blades increases with each row, roughly as shown in figure 6-55. The effects of reheating (increase of gas specific volume) in the flow passages must be taken into account when calculating the gas densities at various sections. Equation (6-136) may be used to estimate the amount of reheat at each row of blades. Also see sample calculation (6-11) and figure 6-60 for additional detail. In the calculations for multirow unsymmetrical blades, the radial heights at the exit side of each row are determined first by equation (6-140). The radial heights at the blade inlets are then made slightly larger, approximately 8 percent, than those at the exit of the preceding row.

Design of Two-Stage, Compounded Impulse and 6-59)

247

Two-Rotor PressureTurbines (figs. 6-10,

6-14

An operational schematic of a typical twostage, two-rotor, pressure-compounded impulse turbine and its velocity diagrams at the mean diameter are shown in figures 6-10 and 6-59. Each stage of a pressure-compounded impulse turbine may be regarded as a single-stage impulse turbine rotating in its own individual housing. Most of the design characteristics of a single-stage turbine are applicable to the individual stages. The gas-spouting velocities Cz and C3, at flow angles a, and a3, of the firstand second-stage nozzles, are designed to be approximately the same. Vz, V 2, V 3, and V 4 represent the relative flow velocities at inlets and outlets of the rotor blades, fiz, fi2, f13, and f14 are the corresponding flow angles for Vz, V2, V 3, and V 4. The second-stage nozzles are designed to receive the gas flow discharged from the first-stage rotor blades at an absolute velocity C2, and to turn it efficiently to a desired angle a 3. Simultaneously, the gases are accelerated to a desired velocity C3, through expansion to a lower pressure. The flow at the outlet of the second rotor has an absolute velocity C4 and a flow angle a4. U is the rotor peripheral speed at the mean effective diameter din. The totalwork performedin the turbineis the sum of thatof the separate stages. These may be designed to divide the load equally (i.e.,the

FIRST

STAGE

FIRST

STAGE

C1 V1

01 "_l

ROTOR NOZZLE

_

_'

'= 2

SECOND

U _

V3

a 3 C3

Figure

6-59.-Velocity

stage, two-rotor, turbine.

STAGE

NOZZLE

diagrams

SECOND ROTOR

of a typical

pressure-compounded

STAGE

two-

impulse

DESIGN OF LIQUID

248

velocity

diagrams

Cl=C

of each

3, C2=C

4, al=a

friction

losses

occurring

passed

on in the

thalpy

and increases

second

stage.

gases not

leaving

The

the lost

carryover

kinetic

gas

ratio

energy

kinetic

stage

rc,

second-stage

nozzles

energy

of the

leaving

from

tance

0.4

to close

between

the

second-stage through should

resulting quire

in equal

drop

may data.

zles

and

With

be used

blades

from

designs

coefficients

(6-122)

and

amount

of reheating.

additional

stage

are

rc

= second-stage

= turbine

ratio

of kinetic

specific specific

through due total in 2

k_

: nozzle

velocity

ent

:nozzle

throat

Sample

Calculation

at constant

F heat

enthalpy

nozzles

can

heat

Btu/lb-deg

ratio

drop the

of the

gases

second-stage

to expansion,

Btu/lb

second-stage

nozzle

coefficient area

coefficient

design

The

available

followfor the

data

2

T2 t = T2 + rc _ gJCp

(6-149)

From sample have been

Y T2t_ )'-1 k_]

(6-150)

(6-11) calculation (6-5), the following obtained for the turbine of the

stage

engine

Turbine

gas

mixture

ratio,

LO2/RP-1

: 0.408

Turbine

gas

specific

heat

at constant

pres-

sure,

p2(

gas

(Ant) 2 =required area,

A-1

P2t:

at second-stage

carryover

=turbine gas

and

nozzles:

C2

at firstft/sec

single-

in the

turbines.

correlations

of second

outlet,

flowing

(6-136)

for the

may be employed for two-stage

velocity

AH2_ 3, =isentropic

or concurrent

established

flow blade

=gas-spouting velocity nozzle exit, ft/sec

y

for noz-

by past

the

gas rotor

pressure,

re-

previous

at second-

psia

energy

of the

enthalpy

pressure

inlet,

C3

Cp

may

in view

at second-

stage

drop

given

turbines

design

stage

°R

psia

=absolute

stages,

equations equations

calculations

leakages

proper

velocity

to estimate

Most stage

the

dis-

the

enthalpy

for each Or, the

C2

pressure

static

at

inlet,

inlet,

nozzle

inlet,

temperature

total

gas

temper-

nozzle

nozzle

nozzle

stage

carryover.

right

approach,

be estimated

experiments,

ing

work

of reheating.

test

the

(stagnation)

static

gas

= turbine

can

axial and

between

of the

a trial-and-error

effects

The

for optimum

determination

stage,

total

gas

= turbine

P2

kinetic

as

P2 t

of the total

well

= turbine

stage

to the

diaphragm

be minimized

The

and

energy

rotor

T 2

turbine.

first

gas

at second-stage

second-stage

inlet

to unity. as

sealing

for the used

ratio

the

= turbine ature °R

as

first-stage

nozzle,

the

T2t

of the

largely

the

where

is

energy

is

ROCKET ENGINES

en-

energy

utilized

by the

vary

stage

a single-stage i.e.,

or

The

additional

available

actually gases

as

the with

identical

first

the first

as

are

4, etc.).

in the stream

Also,

entirely

stage

3, a2=a

PROPELLANT

turbopump.

Cp=0.653

Btu/lb-deg

Turbine

gas

specific

Turbine

gas

constant,

heat

total

ratio,

R = 53.6

Gas total temperature = 1860OR Gas

F

at turbine

pressure

at turbine

y= 1.124 ft/°R inlet,

inlet,

T o

Po = 640

psia C3=kn

gJCpT2t

_

P3

Y

Gas

static

pressure

at turbine

exhaust,

pc=27

psia = kn ,(_-cC22 + 2 gJAH2_

s,

Total

(6-151)

available

gases, (Ant)2

(6-152)

= l /

r27

Y+l _-zT

4/ ,,'LT-qj

content

Btu/lb

Turbine

gas

rate,

Turbine

shaft

Overall

turbine

compounded In addition, forth:

L-

energy

AH = 359 flow

speed,

turbine

_i,_= 92 lb/sec N= 7000

efficiency wheels), the

of the

following

rpm

(when _t= 58.2 design

using

velocity-

percent data

are

set

DESIGN OF TURBOPUMP

Nozzle

aspect

Nozzle

velocity

Nozzle

throat

Nozzle

exit

Rotor

and

ratio

area

kn = 0.96

coefficient,

ent:

coefficient,

stator

blade

(a._) Single-stage,

0.97

compounded

ene = 0.95

velocity

and

stator

blade

exit

area

turbine

length

Cb=l.4 Partition

of rotor

and

coefficient,

at the

tn = tb = 0.05

Solidity

of first

Solidity

of stator

Solidity

of second

blades,

exit

of nozzles

for the

ation

of this

blades

= 1.82

tions

= 1.94

rotor

= 1.67

velocity-compounded,

impulse-type

A-1

stream

stage

engine

reaction of the

and

with stator

compounded,

impulse-type

stage

engine

turbopump,

stage

and

about

turbine with

3 percent

downstream

of the

about

6

of an

reaction

in the

nozzles

of each

3',

oper-

subscripts

processes

listed:

representing

inlet

first

blades;

second

first second

and

rotor

representing

and

condi-

blades;

second

nozzles;

blades;

rotor

rotor

of the

4' =Points

at the

expansion

zles;

rotor

first

second

rotor

the

blades. exit

condi-

rotor

blades;

rotor

blades,

for

1-2,

2-3, nozzles;

in each

blades;

and

1'-1,

stage.

2'-2,

stant

second

3'-3,

4'-4

rotor

blades;

rotor

blades.

and

due in the

stator

along

between

ideal

processes

and

actual

losses

and

and

first

second

rotor rotor

con-

isen-

to friction nozzles,

blades,

processes

= Differences

lines,

expansion

stator

noz-

blades;

of actual

first

pressure

heating

a_ PO

in the

stator

3-4=Path

in the

rotor

process

blades;

blades.

A-1

processes,

'_k_-

diain the

following and

nozzles;

isentropic

0-1,

work

6-60

an ideal isentropic expansion process. 0-1', 1-2', 2-3', 3-4' =Path of an ideal

down-

for the

equal

2',

tropic

0

The

blades;

stator

for

blades

diagrams pressure-

for this

Figure

involved

points

conditions

tions

nozzles.

(b_) Determine the velocity alternate two-stage, two-rotor,

blades

and printwo-rotor,

turbine

turbopump,

in rotor

at the

exit 1',

the velocity diagrams of the single-stage,

percent

turbine. various

stator

blades

(a_) Determine cipal dimensions

the

processes

0, 1, 2, 3, 4 = Points

blades

6-58.

temperature-entropy-enthalpy

gas

the

diagram

in figure

and

in

rotor

the

velocity-

turbine. velocity

shown

gram

denote

in thickness

blades,

stator

is

represents

_b2 =0.95 Chord

two-rotor,

impulse

A representative

coefficient,

kb=0.89 Rotor

249

SYSTEMS

Solution

= 9.7

coefficient, area

PROPELLANT-FEED

re-

blades,

blades

CONSTANT

To A, s

/_PR_LFE

UNES

I

Point :2:

/"

II

_..,"_,'_.,% / p4

>_-

inlet

total bine of the

gram

the

gas

processes

two-rotor, velocity-compounded with small amount of reactions the

nozzles.

drop

available

gases

efficiency

in

about

enthalpy

drop

a single-stage,

rotor

and

stator

impulse turbine downstream o[

drop

in the

6-60.-Temperature-entropy-enthalpy o[

= turbine

= 359

First

of the

energy

tur-

content

Btu/lb

= kn 2 : (0.96) exit=

inlet

psia

enthalpy

= total

"1 "-Nozzle

inlet

2 = 0.92

Rotor

Blade

Inlet

S

Since Figure

pressure = 640

turbine

: turbine

= 1S60°R

isentropic gases

T/n = nozzle Point

temperature

total

pressure

AH = overall

ENTROPY,

total

temperature

= nozzle

-r b==

-._-_/

Inlet

inlet

total

o.

Po

...........

" O"-Nozzle

T O = nozzle

dia-

6 percent AH is blades,

of the

assumed the

overall

to occur isentropic

isentropic in the enthalpy

nozzles

'AH o-1' : A H (1 - 0.06)

-- 359 x 0.94

: 337.5

Btu/lb

250

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

We can write:

Ideally, the efficiency r/n b of a two-rotor, velocity-compounded impulse turbine is a maximum when the turbine velocity ratio

AHo_p = CpT o I 1-(P-21U y_,] \Po]

U

COS

C1From this, zle exit

the gas static

pressure

COS a I U=C1_=3940xCOS

CpToJ

= 640

E1

9._

=

From equation diameter

640 x 0.053 = 33.94

From equation (6-121), the gas velocity at the nozzle exit

(6-122),

(1 - kn 2) C12 _ 0.08 × 15 524 000 =

kn22gJ

Referring at the nozzle pansion

0.92× 64.4 x 778

of reheat C,cosal-U sinai

_ 3940×0.9063940x0.42390=0.622

/31 =31053 '

= 27 Btu/lb

to figure 6-60, the gas temperature exit, following an isentropic ex-

Referring to figure 6-58, the relative velocity at the first rotor blade inlet

gas

flow

C 1 sin a1._3940× sin 25 ° sin fll sin 31°53 ' AHo_,, Cp -1860

TI, ZTo The actual nozzle exit

gas

static

337.5 0.653-1344°R temperature

27 qn---sr T1 = TI' + Cp -- 1344 +0.---_ The gas

rotor mean

From equation (1-130), the relative gas flow angle/3, at the inlet to the first rotor blade can be calculated:

tanfll=Cl

qnr

the turbine

dm =720 U_0_720 × 89___- 29.1 in 7r N 7rx7000

spouting

the amount

(1-129),

psia

CI = kn V2 gJAHo_l, = 0.96 _/16.9 × 106 = 3940 fps From equation in the nozzles

at the mean diam-

=3940 × 0.226 = 890 fps

337 5 ]o.124 0.65-3x-_s60.J

x (0.722)

speed

425 °

1.124

=640×

1

at the nozFrom this, the peripheral eter of the rotor

Pl =Po

a

4

density

at the nozzle

at the

= 1385° R exit

P, 144 33.94x 144 Pl =-_-T x--_ -1385.4 × 53.6 = 0.0658

lb/ft 3

We will use an angle a_ of 25 ° for the spouting-gas-flow direction at the nozzle exit.

_ 3940 x 0.423 O. 528 Point Inlet

"2"-First

Rotor

Blade

Exit=

3156 fps

Stator

Blade

Assume that the given 6 percent reaction downstream of the nozzles is equally divided between the two rotors and the stator. Then the isentropic enthalpy drop in the first rotor blade can be approximated as

AH,_2, =P-_

x 359 = 7.18

Btu/lb

Using equation (6-135), the relative gas velocity at the exit of the first rotor blades

flow

DESIGN OF TURBOPUMPPROPELLANT-FEED SYSTEMS

V 2 = V'kb 2V l2 + 2 gJTlnAH l_2, = V/(0.89 × 3156) 2 + 64.4 × 778 × 0.92 × 7.18 = 2866

V 2 sin/32 tan a 2 = V cos /32 - U a 2 =35°15

fps

From equation (6-136), in the first rotor blades,

the amount

of reheat

251

2866×sin 25 ° 2866 × cos 25 °- 890 = 0.707

'

The absolute blade exit

flow velocity

at the first

rotor

c2-V

2 sin /32 _ 2866 × sin 252 1210 2080fps sin a 2 sin 35o15 ' =0.57---_=

(3156)2 = [1- (0-89)2] x 64.4 x 778 _(1-099)×7.18 • ,

Point Blade

"3"-Stator Inlet

= 41.975

The isentropic blades

V2 2

1

qbrl =(1 - kb )_-_+

The static blade exit

(1 - r/n) AHI_ 2,

Btu/lb

gas pressure

at the first

Exit : Second

enthalpy

Rotor

drop in the stator

AH2_ 3, = AH 1-2' = 7.18 Btu/lb

rotor

gas

Y

F A.,

Blade

Analogous to equation (6-135), the absolute flow velocity at the stator blade inlets C a = x/kb2C22 +2gJ_nAH2_

a,

= X/(0.89 × 2080) 2 + 64.4 × 778 × 0.92 × 7.18 =33.94x

1

= 1938 fps

0.653×1385

= 33.94 x 0.93 = 31.6 psia

The gas static temperature first rotor blade row following expansion

Reheat

in the stator ,

at the exit of the an isentropic

2_

qbs =(l-Kb

blades 622

)2--_+ (1- rln) AH2_ 2 (Analogous

= [i- (o.89) 2] × (2080)2 T 2, = T 1 - AH,_2,/C p = 1385 - 7.18/0.653=

64.4 × 778

1374 ° R

to eq. (6-136))

+ (1 - 0.92) × 7.18

= 18.53 Btu/lb The actual static rotor blade row exit

gas temperature

The static exits

qbr2 41.975 --1374-+ T2 = T2' + Cp 0.653 Gas density

at the first

at the first gas pressure

- 1438 ° R

rotor blade

144p2 144x31.6 P2 = RT 2 - _ 1--_-_-8 = 0.059

exit

at the stator

Y p3 =p2 [1 -CDT2 _H2-21 j Y-1 =31.6x =29.42

7.18 _9.06 E1 0.653×1438J

psia

lb/ft 3 Gas static temperature at the stator exits following an isentropic expansion

We use an angle/32 of flow direction at the first symmetrical blades). The a 2 at the first rotor blade from

blade

25 ° for the relative gas rotor blade exits (unabsolute flow angle exits can be calculated

T3, = T2 -AH2_2/Cp=

1438-7.18/0.653=

blade

1427 ° R

Actual staticgas temperatureat the stator blade exits

252

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

qbs 18.53 T3 = T3'+-_--= 1427 + x-_--_. = 1456° R

Gas density

at the stator

blade

}r

exit = 29.42 x

144 P3 _ 144 x 29.42 .= 0.0544 P3 = R T 3 53.6x1456

lb/ft 3

We use an angle a3 of 35 ° for the absolute gas flow direction at the stator blade exit (a s Ta2). The relative flow angle t33 at the stator blade exit can be calculated from

tan/33-C3

C 3 sin a s 1938x 0.574 cos a3- U=1938x0.819-890

= 1.596

f13 = 57°56' The relative

flow velocity

at the stator

blade

= 27.46

is slightly higher sure (underexpansion), effects.

The gas static rotor blade sion

T4' --T3

V3 = C s sin a 3 _ 1938 × 0.574 - 1312 fps sin /33 0,847 "4"-Second

Rotor Blade

The isentropic rotor blades

enthalpy

exits

1 - 0.6537.1s x 1456] -1_°_

psia

P4

exit

Point

=v

r

> 27 psia

than the turbine exit presbecause of the reheating

temperature following

T4

at the second

an isentropic

expan-

7.18 - 1445 Btu/lb -AH3__,/C p = 1456--0.653 -

The actual gas static ond rotor blade exit

Exit

(Pc)

--

T4 '

drop in the second

temperature

qbr2 1445 ' 7.73 + C--_---= +_=

at the sec-

Gas density

at the second

1457 ° R

rotor blade

exits

AH___,= AHa_f --7.18 Btu/Ib The relative gas flow velocity rotor blade exit

g

4

= \/kb

2 V3

2

+

2

at the second

gJ77nAHa.. 4,

= 1306 fps of reheat

in the second

(unsymmetrical blades). The absolute flow angle u 4 at the second rotor blade exits can be calculated from

rotor tan a 4 =y4

V 4 sin /34 1306 × 0.695 cos /34- U=1306x0-719-890

a 4 =86°55 ,,y

= [1 - (0.89):]

flow velocity

at the second

rotor

(1312) 2 , ,. x 64--A x 778" Ll - 0.92) x 7.18 C4 =V 4 sin/34_ sin a 4 Nozzle

exit

'

The absolute blade exits

(1- 9n) AH_-4'

Btu/sec

Gas static

-18"5

2

qbr2 = (1 - kb2)-_gj+

= 7.73

lb/ft a

We use an angle/34 of 44 ° for the relative gas flow direction at the second rotor blade exits

= V'(0.89 × 1312) 2 + 64.4 x 778x 0.92 x 7.18

The amount blades

144 P4 _ 144 x 27.46 = 0.0506 P4 - RT 4 53.6 x 1457

pressure

at the second

rotor blade

1306x0'695=908 0.9985

fps

Dimensions

From equation zle throat area

(6-123),

the required

total

noz-

DESIGN OF TURBOPUMP

wt

PROPELLANT-FEED

First

Ant = /

Rotor

The

y+l

253

SYSTEMS

Blade

pitch

Dimensions

or blade

(at

dm)

spacing

r 2 i?-_

gyL iJ

Blade Pbrl =

chord

length

Blade

Cb

solidity

1.4

= 0.769

- 1.82

in

92 From 0.97 x 640

= 13.22

We use nozzle

x 1.124(0.94) 53.6 x 1860

ll/32.2

equation

in2

a radial

throat.

Zbrl height

Thus

the

hnt of 1.5 nozzle

inches

width

atthe

at the

Allow inlet

throat

aspect

number

ratio

- 9.7 - 0.1548

tive

Ant

Pitch

-

bnthnt

0.1548×

_

1.5

57

2 ° between

spouting-gas

From

The

a blade

nozzle angle

-2=25-2=23

(6-125),

the

exit

angle

al;

thus

+2°7'

=3't°

equal

toexit

rela-

=_2

= Pbrl

sin

=25°

radial

height

1.64 x 1.08=

width

Oblrl

at the

at the

- tb = 0.769

inlet

1.77

in

inlet

x 0.559-

0.05

= 0.379

in

From equation blade exit area total

(6-138),

the

required

total

noz144 x 92 --

p2V2Eb2

0.059

x 2866

x 0.95

82.5

in 2

144 x 92

plClene

0.0658×3940×0.95

53'75in2

Combining obtain

equations

radial

and

On and

° required

Ob2rl

144 w_

144 w_

height

(6-125) and

width

and at the

(6-126),

we

nozzle

exit:

the

equations blade

radial

sin

hne = rrd m sin On - Zntn - rrx29.1x

0.391-

and

(1-140a),

at the

we

exit

Ob2rl-Zbtb

53.75

Ane

(1-139) height

Ab2rl hb2rl-_dm

= 1.64

= 31°53'

(1 x 0.08)= passage

area,

Combining

angle _)b_ri

thus

/_2

=

obtain

-119

blade

fl_;

bladeangle

blade

Ab2rl

Ane

inlet

angle +2°7'

angle

hb,rl =hne

flow

equation

exit

--_1

We select

spacing

On=a1

zle

0.769

flow

exit

flow

bblrl nozzle

,Tx 29.1

Ob2rl

Pn - ,din _ _ x 29.1 = 1.604 in zn 57 We allow

,,dm

13.22

or nozzle

of blades

in

of nozzles

Z/l=

number

1.5 Make

The

the

= Pbr 1

2°7 ' between

relative Obxrl

hnt bnt =Nozzle

(6-140a),

lTns

82.5

57 x 0.05

x 29.1

x 0.423

- 119 x 0.05

= 2.52

in

in The

blade

passage

width

at the

exit

Ane

zn bne = hne

bb2rl

53.75 - 57 x 1.6_-

0.576

in

= Pbrl

sin

= 0.291

in

Ob2rl

- to = 0.769

x 0,443

- 0.05

DESIGN OF LIQUID

254

The

mean

PROPELLANT

blade radial height total

hbr,

-

1.77+2.52 2

ROCKET ENGINES

Using equation (6-138), blade exit area

- 2.145 in

144 w_

required

0.0544

x 1938

x 0.95

= 132.5

in s

(6-139)

and

(6-140a),

we

a tapered blade with shroud, and that

it is subject to approximately

the same

tensile

stresses from centrifugal forces, as would uniform blade without shroud. be made

the

144 x 92

Ab2 s = P3 C3eb2 Assume

we obtain

of Timken

lb/in _. Cheek

The

be a

Combining calculate

equations

the

blade

radial

height

at the

exit

blades shall

alloy, with a density pb =0.3

Ab2s

the centrifugal tensile stresses

sin

hb2s=rtdm

Ob2s-Zbstb

at the root section using equation (6-141). 132.5 x 0.574 - 127x

- _x 29.1 Scr I = O.O0045721pbhbrldm

The 0.0004572 × _-0_-X, × 2.145 × 29. i X

blade

passage

bb2 s = Pbs

Second

From

chord length Cb solidity

ndm_

Allowing 0b, s and

Obls=a2

of blades

absolute

inlet flow

angle

blade

exit

angle

a 2

length

(6-140a),

2o4 ' between

and

the

Cb

solidity

equation

Allow 0blr2

x 0.574-

0.05

Dimensions

chord

Blade

From blades

inlet

1.4 1.67

the

0.888

number

in

of the

_ v×29.1=i09 0.838

the

relative

inlet flow

blade

angle

angle

f13;

thus

Oblr2 =f13 +2_4' = 57_56' +2°4' =60o

+2024 `= 34°36 '+2o24 ' =37 °

8b2s=a3

at the

spacing

vdm Zbr2 - Pbr2

We hold exit.blade angle 0b2s equal to exit

We make

the exit blade angle 062r2 equal to

the exit relative flow angle /94 Ob2r2 =/_4 =44°

=35 °

equation (6-149), blade radial height at

From

equation (6-149), the blade radial height

at the inlet is

the inlet

hb,s= The

=

- 127

absolute flow angle a3:

From

Pbr2

0.721

2024 ' between inlet

Blade

Blade

- 1.94 - 0.721 in

_× 29.1

width

062 s - Ib = 0.721

or blade

1.4

equation (6-140a), the number

Zbs = Pbs

in

in

Rotor

Pitch

Pitch or blade spacing

Blade

x sin

= 0.364

Stator Blade Dimensions

Pbs -

-2.87

(7000)2

= 13 050 psi

Blade

0.05

N2

l.08×2.52=2.72

blade passage

bbls= Pbs =0.384

width at the inlet

sin Ob,s-tb=O.721xO.602-O.05 in

hbir

in The

blade

2 = 1.08× 2.87= 3.10 in passage

bb 1r2 = Pbr2 sin = 0.677

in

width

at the

Ob l r2 - tb = 0.838

inlet x 0.866-

0.05

DESIGN OF TURBOPUMP

From blade

equation

exit

(6-138),

the

-P4

_:

total

the

A-1

144 × 92

V4eb2

Combining obtain

-

x 1306

0.0506

equations blade

x 0.95 = 211

(1-139)

radial

and

height

in 2

(1-140a),

at the

we

exit

Stage

Rotor, Design

hbsr2-_d

m sin

velocity

figure

fps; 119×

0.05

diagrams

exit

passage

sin

= 0.533

in

mean

blade

Obsr:

=3"66

in

V 2=2866

x 0.695-

_4=44_;

C4:908fps;

height

3.10+ 2

root

the

- 3.38

section

centrifugal using

tensile

equation

Nozzle

ratio

(at

= 9.7;

=0.1548

in;

rotor

2

× 29.1

Efficiencies

From bined

equations

nozzle

and

blade

and

(6-147),

the

com-

efficiency

in;

t)blrl

=34°;

hb2rl

=2.52

=0.291

blade

a 1 +C:

cos

bb28=0.364

Second

rotor

Pbr2

tinb --

a 3 +C 4 cos

a4)

890 (3940

× 0.906

+ 2080

x 0.817

+ 1938

x 0.819

32.2

x 778x

359

equation

the

turbine

machine

(at

in;

din):

Zbr2 = 109;

in;

Oblr2

=600;

0b2t2

hbsb2

=3.66

in;

compounded grams, following

(6-148),

dimensions Cb = 1.4

in;

=0.533

see

fig.

prior

obtained. rc = 0.91.

=44°;

bblrs=0.677

in

two-rotor,

impulse

turbine.

equal-work, (For

pressure-

velocity

dia-

6-59.) trial-and-error

isentropic

(approximately) From

bbls=0.384

=0.838

From

= 0. 683

efficiency

blade

(_b) Two-stage, + 908 × 0.055)

Zbs = 1.27; 0b28=35°;

in

= 1.67;

bbsr2

=25¢;

bblrl=0.379

hblr 2 =3.10 in;

gJAH

in;

0bls=37°; bb:s=2.87in;

in;

a2

÷C 3 cos

Obsri in;

(at din):

Cb = 1.4

in;

din):

in

dimensions

= 1.94;

Solidity U(C 1 cos

Zbr_=llg;

in;

hbls=2.72in; (6-146)

(at

Cb=l.4in;

= l.77 bbsrl

in; bnt

dimensions

=0.769

Solidity

psi

Pn = 1.604

nne=l.64in;

bblrl in;

x (7000):

din):

bne=0.576in

blade

Pbs=0.721 Turbine

Btu/lb

Pbrl

Stator = 20 550

Btu/lb

Btu/lb

z n = 57;

hnt=l.5in;

Solidity=l.82;

3.38

AHI_ 2, = 7.18

AHs. _, = 7.18

in

0n=23°;

at the

(6-141)

x_x

fps

drops:

dimensions

First

= 0.0004572

a4 =86°55';

V4=1306

efficiencies:

dm= 29.1

in

stress

Scr 2 = O.O0045721pbhbrsdmN

fl3=57°56';

fps;

rotor, AH3. _, = 7.18 5H = 359 Btu/lb

Aspect Check

V1 C 2=2080

7/t=58.2%; ;7n=92%; _7nb=68.3%; qm=85.2% Mean diameter of nozzles and blades:

3.66

hbr2 -

a3=35°;

blades,

blades,

Second Total radial

din,

AHo_ I,= 337.5 gtu/lb

rotor

Working The

diameter

fl_=25:;

V 3 =1312

enthalpy

Stator

0.05

fps;

fps;

Nozzles,

width

- tb = 0.838

Type)

C_=3940fps;

a:=35°15';

C 3 =1938

First = Pbr2

at mean

fl_=31°53';

Isentropic blade

Two-

Impulse

6-58.

=3156fps;

211 x 0.695-

(Single-Stage,

U = 890:

Obsr2-Zbtb

=_x29.1

Turbine

Summary

For see

Engine

a 1=25°;

bbsr2

0.582

Velocity-Compounded

Absr2

The

255

SYSTF.j_,S

area

144 #t Ab2r2

required

PROPELLANT-FEED

equal We assume

enthalpy work a stage

calculations, drops

the

resulting

for each carryover

stage

in were

ratio

256

DESIGN OF LIQUID PROPELLANT

First-stage

nozzles:

AHo_I, = 50%;

AH = 0.5 × 359=

179.5

ROCKET ENGINES

Point

"2"-First-Stage

Stage

Nozzle

Btu/lb

From

equation

velocity First-stage

rotor

(6-135),

at the

V 2 : \/kb AH : 0.03

× 359 = 10.75

Btu/lb

AH2_ 3, = 44%;

= 1736

AH = 0.44

rotor

AH3_4, = 3%;

× 359 = 158

We chose

Btu/lb

blades:

absolute lated as

AH = 0.03

× 359 = 10.75

" O'-First-Stage

Nozzle

tan

Inlet

a s-

a relative

gas

flow

a2= 86040

"1 "-Firs$-Stage equation

Nozzle (6-121),

first-stage

= 2880

Exit

the

nozzle

C 1 = kn \/2 gJAHo_

flow

= Rotor

Using /3_ at the calculated

a l.

For

× 223.8

optimum

--

equation

flow

angle

blades.

can

then

The be calcu-

1736 1736

×0.616

x 0.788-

1308

17.25

0.906×2880 2

sin

rotor

the blade

sin

al

Blade

inlet

equation

gas

Nozzle

(6-151),

the

can

be

stage

V'0.91

: 2880

fps

C 3 = C_,

velocity stage,

C 4:C

+ 2 gJAH2_

= 0.96

Since

flow

Exit

= Second

second-stage

noz-

× (1070)

the

a3:a fps;

2=1070fps;

2 + 64.4 × 778 x 158

remainder

diagram i.e.,

3'

is

the

I =25°;

of the same

as

secondthat

of the

J33 =]31 =43°8';

a4:a2=86°40';

/_4=/_2=38°;

V 4=V_=1736fps.

- 1308 = 0.936

'

The relative blade inlet

fps

velocity

C a = k n \/rcC22

From 131 =43°8

- 1070

Inlet

gas-spouting

1308fps

relative

× 0.906

first-

tile

2880×0.423 U - 2880

at the

0.998

"3"-Second-Stage

first

al-

velocity

1736×0.616

a2

Rotor

diameter

_

flow

exits

Point

zle

spouting-gas

mean

gas

× vJl-7-9_5.5

efficiency,

rotor

(6-130),

first-stage as

/31 = C1 cos

Yl

a2,

f_2

blade

C2 :

exit

f = 0.96

at the

cosalC1 2

C,

rotor

× 10.75

'

V3=Vl=1784

tan

gas rotor

angle,

V 2 sinfl2

gas-spoutingveloc-

of 25 ° for the

speed

--

flow

Blade

fps

a value

angle

peripheral

exit

/32- U

absolute rotor

From We use

× 778 × 0.92

first-stage

i,,2 cos

The

ity at the

exit

psia

stage

From

gas

blade

Btu/lb

T O-- 1860oR

Point Inlet

relative

rotor

fps

V 2 sin

Po = 640

Exit=Second-

2V12 + 2 gJT?n'_H1_ 2,

/3 2 --38 ° for the

Point

the

first-stage

: V_(0.89 × 1784) 2 + 64.4

nozzles:

Second-stage

Blade

blades:

AH__ 2, = 3%; Second-stage

Rotor

Inlet

equation

(6-129),

the

turbine

rotor

mean

diameter gas

C1 sinai sin 131

flow

velocity

at first-stage dm=720U-720x1308 _N _r × 7000

2880×0.423 0.683

1784

From

fps and

blade

equation efficiency

(6-147),

the

42.7

combined

in

nozzle

257

DESIGN OF TURBOPUMP PROPELLANT-FEED SYSTEMS

result

U(C x cos a_ +C 2 cos a 2 +C 3 cos a3+C 4 cos a4) r/nb =

g JAil

right angles those which loads).

1308 (2880 × 0.906 + 1070 x 0.058 + 2880 × 0.906 + 1070 × 0.058) 32.2 × 778 × 359 =0.78

sult The turbine machine efficiency is assumed be the same as that used in design (a._):

to

qm = 0.852 From equation efficiency

(6-148),

the overall

turbine

r/¢ = rlnbr/m = 0.78 x 0.852 = 0.664 A-I Stage Engine Alternate Turbine Design Summary (Two-Stage, Two-Rotor, PressureCompounded, Impulse Type) For velocity diagrams at mean diameter din, see figure 6-59. U = 1308 fps: a1=25°; _1=43°8 '' C1=2880fps; V 1=1784 fps; a_=86°40'; _2=38°; C2=1070 fps; V 2 = 1736 fps; a3 =25°; f13 =43°8'; C3=2880 fps; Va=1784 fps; a4=86°40'; _4=38°; C4=1070fps; V 4=1736fps Isentropic enthalpy drops:

6.6 DESIGN SEALS, Turbopump

a relatively

OF TURBOPUMP AND GEARS Bearing

large

or on the

These forces may those which act at

axis (radial to the shaft

forces), and axis (thrus:

Radial loads on turbopump bearings may refrom one or more of the following sources: (1) Weights of parts such as shafts, pump impellers, turbinerotors, gears (2) Centrifugal forces du_:, to unbalance of these rotating parts (3) Forces due to inertia, resulting from rapid acceleration (4) Resultant radial forces on the impeller due to nonuniform pressure distribution in the discharge volute of the pump (5) Tangential or torque forces induced by the gears on turbopump bearings may remore of the following sources: rotating parts mounted on a shaft

For the turbopumps of liquid rocket engines, high-speed ball and roller bearings are used almost exclusively. A typical two-bearing design is shown in figure 6-7. A ball bearing carries both radial and thrust loads. It is paired with a roller bearing which carries only radial loads, however, of a higher magnitude. A typical three-bearing arrangement is shown in figure 6-63. The shaft radial loads are carried by a single roller bearing at the turbine end and by a roller and a ball bearing on the pump side. The ball bearing also absorbs the thrust loads. As a rule, the shaft thrust loads in a turbopump are carried by a single or dual bearing located at one end of the shaft. Thus loads from thermal expansion or contraction of the shaft are avoided. Bearing design data with regard to loadcarrying capacity, operating speed, and service life are usually furnished by the manufacturers. The useful life of a bearing is dependent upon its speed and load, and may be expressed by the correlation:

that dm is

BEARINGS,

Design

A turbopump shaft is supported by two or more bearings. The loads on the bearings are the

-

act on the shaft

(2) Unbalanced axial thrust of the pumps (3) Resultant axial thrust on the turbine rotor blades

Working efficiencies: r/t= 66.4%; r/n = 92%; rlnb= 78%; _?m= 85.2% Mean diameter of nozzles and blades:

of design (a). However, required (weight, size).

which

to the shaft act parallel

Thrust loads sult from one or (1) Weight of vertical

First-stage nozzles, AH__,, = 179.5 Btu/lb First-stage rotor blades, AH,_ 2, = 10.75 Btu/lb Second-stage nozzles, AH2_2 -- 158 Btu/lb Second-stage rotor blades, AHz._4, = 10.75 Btu/lb

dm= 42.7 in Comment: The overall efficiency of the pressure compounded turbine is higher than

of forces

parts supported by the shaft. be divided into two classes:

lira

_

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

258

{ Rated speed, rpm Life, hours = gb \Actu-a--1o--p-er_p-_'-d,

._ rpm/

[ Rated capacity, ib ,_4 ×_Actual working load, lt)] (6-153) where

Kb = design factor usually manufacturer.

furnished

by

If a bearing is subjected to both thrust and radial loads, the two can be combined into a single

equivalent

radial

load:

P = R + xA

(6-154)

where P = equivalent radial load used for bearing selection, lb R = actual radial load, lb A = actual thrust load, lb x = design coefficient usually furnished by manufacturer Rocket turbopump bearings quite commonly are cooled and lubricated by the propellants pumped. They are usually operated at very high "DN" values, a parameter which is the product of the bearing bore D (millimeters), and the bearing rotative speed N (rpm). Propellant lubrication has the advantage of eliminating an additional lubricant supply system, and of simplifying bearing sealing problems. The following are important design considerations for propellant-lubricated bearings: (1) Characteristics of the propellants, such as thermal stability, operating temperature, chemical inertness, viscosity. (2) Compatibility of the bearing materials with the propellants. The application of certain high-strength alloys is sometimes limited by the propellants used. The "DN" rating is convenient when selecting high-speed ball or roller bearings. As the rotatire speed of a bearing increases, contact fatigue of the outer race caused by centrifugal loads of the balls or rollers may cause failures. In addition, bearing contact speeds will result in nonrolling phenomena with attendant failures caused by overheating. Through proper selection of the bearing geometry, these problems can be minimized, and the DN rating increased. Note that for a given horsepower rating, the shaft size based on allowable stress does not decrease

proportionally with the increase of shaft design speed. Thus the required bearing DN value rapidly increases for high-speed turbopumps. As a result, especially for liquid hydrogen application, the turbopump rpm is often determined by the DN limits of the bearings. A most important bearing design consideration is the expected operating life of the rocket engine. The bearings must have adequate statistical probability of conforming with this requirement. A generally accepted life rating for ball and roller bearings is the "B-10 life." The term denotes the operating life (hours) of a _population" of bearings at a given load and speed, at the expiration of which statistically 10 percent of them will have failed. Of course, in actual rocket engine operation, component reliability must be much higher. Bearing life at a given load and speed varies inversely with reliability. For instance, the B-1 life (99 percent reliability) is one-tenth of the B-10 life (90 percent reliability), or one-fiftieth of the B-50 life (50 percent reliability). Therefore, turbopump bearings are generally designed for a B-10 life of at least 100 hours. This corresponds to a B-1 life of 10 hours, or a B-0.1 life (99.9 percent reliability) of 1 hour, the latter by order of magnitude being the life the bearing most likely will actually see. For critical applications, an even higher life rating may be selected. Figure 6-61 presents the centrifugal load DN limits in terms of 10, 100, and 1000 hours of B-10 life for a typical ball bearing design (extralight series). The stress-limiting DN values of roller bearings are much higher than for ball bearings; however, it is extremely difficult to control the temperature rise in a roller bearing, if the DN value is above 1.5 × 106, due to excessive cage slip. Generally, rocket turbopump bearings have been successfully operated at DN values up to 1.5 × 106. Limited test information indicates possible

satisfactory

Dynamic

Seal Design

operation

at 2.0 x 106 DN.

The principal dynamic, i.e., rotating seal types used in liquid rocket turbopumps are the labyrinth, face-riding, and shaft-riding seals. Satisfactory seal operation depends upon good design which considers many factors, including

DESIGN OF TURBOPUMP

PROPELLANT-FEED

_.:1

I

I I II I

[

I

¢ol

I

J

I

I ]

II -

I

I

It l

•

[

I

I

_.ou_e

&o i

o

i J

$O MRS

i

i

J

[

=,

BALL

BEA

NG

C

. TRt N

IFUGAL

II

r

.55

J

I

_ _ ._ I I I I

",., .c__..c- 2:.':.

1

LIMIT

IOUTE_ =.cE _*FO*=,T. T--SPEC, .58

../---TST'0,

\-T-_

_.o#-------------J_/--\--4_,_.

1

LOAD

(EXTRA L,_,TSE=,ESS,ZEI !

[ '

I

259

SYSTEMS

...

--_-\-q----

-_ _ ---

' _.3u-=_.

't,

,

I .....

o. Io

ZO

3o

40

50

60

70

8090

_

200

_0

400

=X>o

BEARING BORE SIZE, D mrn Figure

fluid

pressure

surges,

contraction between

ties

of the

free

operation

of the

ence

directly

which these seal.

As

The

function

fluid

p

= density

pres-

Cs

= seal

frictionand

Any

improper

of the leakage

6-62,

seal.

The

sealing is

forced

Labyrinth

on

of pump seals

influ-

seals

leakage

to follow

a devious

entirely,

tion

and

wear. seal

is

but level

The can

amount be

seal

throttled is

rather

not

turbine

of leakage

is

Aps/p

are

path.

=leakage

Ac

=seal

rate, clearance

leak-

turn,

of friccorrela-

(6-155)

a

seal

angle 6-62,

welded is

casing.

to the

follow

axial

and

Sometimes

Aps = pressure lb/in 2

differential

across

seal the

seal,

seal the

and

flexibility

seal

is

(fig.

The

segments

held

against

to the and

permitting

movement used washer,

6-62)

is in

housing

provides face,

As

washer

bellows,

seal

retaining

are

seal

The

floating

faces

of rotation.

floatin£

a seal

segments. shaft

seal

mating

sealed

seal

is between

statically

angular

shaft-riding

housing,

axis

contact

a lip

sealing

The

bellows.

bellows

force

The in 2

between

of a floating

the

and

spring age.

rings

rotating

contact

to a stationary

secured The

the

ring. to the

to a metal

is

wearing

for the diaphragm

faces

in figure

which

for the as

rubbing

a shoulder

at a right

attached

in3/sec area,

3 established

to tim sealing

with a spring-loaded a bellows.

Qe

used well

through

and

shown

to prevent

where

lb/in

stages.

precision-lapped

many

tion Qe=CsAcV'24g

the

washer

through

by the

as

In a face-riding-type

to pass

to reduce

at a minimum

estimated

seal

tending

are

impellers,

attached

two

operation

labyrinth

fluid

interface

of a labyrinth

to a reasonable

the

fluid

coefficient,

experimentally

subsequently

cause

in figure

age

labyrinth

can

and

DN limits.

accomplished

the and

and

surfaces.

load

veloci-

parts,

or indirectly

factors

shown

times

smooth sealing

sealing

a clearance-type through

rubbing

surfaces,

of internal

centrifugal

contact

surfaces,

sealing

bearing

expansion

components,

sealing

squareness alters of the

vibration,

of sealing

sure

6-61.-Bali

it to

without

leak-

in conjunction instead

consists

plate,

and

form

a ring

it by garter

of

of a several around springs.

260

DESIGN OF LIQUID PROPELLANT

ROCKET

ENGINES

-PUMP CASING WEARING RING

....,..\\_

HOUSING

/-- PUMP IMPELLER

___

-_-_-

_SHAFT

CC SHAFT LABYRINTH

SEALS

_

_-TURBOPUMP FLOATING .\'..x>,_" SEAL WASHER_\_\_/ST

,.AL

RETAINING

PLATE /-_

CASING

TURBOPUMP

/CASING

P- SE;TL HO_SING ICSE L

____GARTER

SPR,NGS

_HER_/SEAL SHOULDER

RING_

HOUSING

BELLOWS ....._/,/_ _

_

___cu-_L_L

FACE-RIDING

- --

SEAL

SHAFT-

Figure

Thus

self-adjusted

plished

dynamic

between

ment

inside

forced

shaft

diameter.

against

a static

to occupy

less

A wide floating seal

variety

is

hardened

The 300

assures such

face

rubbing

sealing

interpropellant

tend seals. for

speed

smooth

two

in series.

for critical

held

as

not

gears must

or

sistance

The

gear

pumps

(fig.

tween

turbine,

also

sometimes

trains 6-16)

used afford

pumps between

in liquid speed

and

turbo-

differentials

accessory a pump

rocket

very

impeller

and and

a

often

internally

gears

designer,

steel,

the

with

tooth

surfaces

or induction surface

very

should

process.

tolerances under

re-

usually

tooth

by a grinding

be held

high are

carburizing

dimensional

on bear-

therefore, and

gears

the

most

in turbopump

strength

case

are thrust

Turbopump

finished must

are

of turboclose

control

manufacturing. improve

certain

practices

The

If possible, and

sections

speeds

or

are

cross

minimize

tooth

by either

gears

To ity,

high. high

webs

are

Spur

and

of high-alloy

during

be-

drives,

results.

to wear.

pump

hubs

loads

hardening.

Design

The

oil,

in an aluminum

further,

as

Tooth

Materials Gear

possible

geom-

operation,

with

housed

weight

they

be accurately Turbopump

pumped.

since

hardened

applications

being

used,

are

and

propellants,

During

lubricated

usually

hub.

achieve

made

seals.

as

for best

ings.

factors. and

minimize

thin

rim and

lines This

are

To

widely

should

between

gears

splined

be

or purge

The

arrangement transmitted,

other

cooled

propellant

casing.

the

rubbing must

to a very

vent

installed

the

segments.

The

lapped

cavities

seals

positive as

ring

or shaft

Frequently, to the

dynamic

seal

ring

with

Gear power

and are

SEAL

types.

upon

ratio,

thus

available

frequently. and

seal fps.

connected

more

most

or plated,

exceed are

and

speed

RIDING

inducer.

gears

seals

seal

depend

the

face-riding is

are

etry

housing

washer,

of materials

on shoulder

finish.

the

seg-

segments seal

dynamic

low-speed

and

Shaft-riding than

washers

used

faces

the of the

a spring

seal. space

seal

Carbon

and

turbopump

is accom-

diameter

surface

plate

providing

sealing

Axially,

a flat

by a retaining

6-62.-Principal

outside

SEGMENTS

.....

_ SHAFT

gear

life

modifications can

be applied.

and load-carrying to standard Pinions

stabildesign

are

frequently

at

DESIGN OF TURBOPUMP

made

with long addendum

addendum

PROPELLANT.FEED

considerations

and gears with short

to adjust tip-sliding velocities and to

strengthen the pinion. thicknesses

Furthermore,

pinion tooth

are often increased, at the expense

of gear tooth thickness.

High pressure angles

as high as 22_ °,25 _j,or 271/: ° may be applied to reduce contact stresses on filetooth surface and to increase the width of the tooth at the base,

to compensate

tips

from

cutting

the

The

DESIGN LAYOUT ASSEMBLIES

OF

is

layout aging

a list

Figure 6-63 presents the design layout of the A-1 stage engine turbopump assembly. Logical packaging and arranging of the basic mechanical elements of tim turbopump are among the

For

of important

Positive

systems

pack-

integrity interpropellant

sealing

for thermal

(5) Ease of development (6) Ease of assembly (7) Ease of manufacturing Considerable experience in turbopump

turbopump

plumbing

(4) Compensation contraction

TURBOPUMP

layout.

with engine

and

(2) Structural

part.

the

considerations:

(1) Compatibility

(3)

6.7

in preparing

following

for bending and to keep the mating

261

instance, one of the more important criteria which influences the selection or arranging of the turbopump mechanical elements is the ease of development. Standard or proven mechanical detail should be extensively adopted in the layouts. design

Involute-profile modifications are often also made

SYSTEMS

design

layout

expansion

and

and skill are required work

for best

results.

iNLET PUMP PUMP VOLUTE AND DISCHARGE

i _-FUEL

PUMP

VOLUTE AND DISCHARGE MANIFOLD i

FUEL

PUMP

-TURBINE

NOZZLES ROW

ROTATING

BLADES IMPELLER-_

OXIDIZER PUMP

ROTATION

/

OXIDIZER PUMP

OXIDIZER PUMP IMPELLER

ROTATING BLADES SEAL

INDUCER

HYDRAULIC

3TATIONARY PUMP

IMPELLER

BLAOES

AUXILIARY POWER P1CKUP SHAFT

Figure

6-63.-Assembly

design

layout

of the

hypothetical

A-1

stage

engine

turbopump.

Chapter Design 7.1

CONTROL The

Controls

of

METHODS

opening

foremost design requirements for any

and

deviations pellant

control system ability. Two able:

closing from

flow

the

the

and

Valves

propellant

valves.

design

rates,

VII

mixture

such

as

from

Minor

ratio

or pro-

fabrication

are accuracy, stability, and relibasic control methods

open-loop

are avail-

(no feedback) and closed-loop

tolerances

of engine

beforehand

by insertion

orifices (feedback) control systems.

Both have

into

the

components,

are

corrected

of accurately

propellant

sized

flow

lines

(also

see

to effect

found the

desired

pressure

drops

ch.

II).

The

wide application in liquid propellant rocket proextent pulsion systems.

Open-loop

to those systems

which

of correction

to operate

preflight has

calibration

the

conditions.

Most other applica-

tions require one of the many can be constructed with which

and

is unable

tions

during

models

such as gain factors and stability of a then

system

such

characteristics of the

being controlled, and by allowable Once

elements

the method

time-

is determined, the basic

for the proposed

system must

lected, such as type of components

locks. their

It will be influenced by the required

accuracy, the dynamic lags.

be se-

to compensate

the specific control, which specific application.

of an open-loop

control

for engine

and

of the power

lant

valves

ators

are

with

fuel

and

the

relative

pets,

development

all depend

start

the

aid

are

preferred

stop

of inter-

for

other

of many

positions

with

work.

However,

some

work with attendant redesign will

requirements

Control

off command

means

control

means, devices. is

calibrated

propellant

achieved valve also

between by adjusting

gates

mechanical

can

(electric,

or pop-

linkage.

be

such

is accomplished as

orifices,

A typical an engine to a fixed

flows

hydraulic,

engines, valves

interlock

furnished

by

or pneumatic).

sequencing

and

by the

quence

designs. engine,

ignition

between

system

combination

and

are

propellant

controlled

by

and

example

set

A typical

for which their

the

interlocks

in chapter

Closed-Loop

system,

control

gener-

operated

main

is often

of various example

is the

start

and

stop

were

described

A-1

sein

of perfecting a

system.

this

to the

are

sequencing

is of the

respect

for propel-

or gas

and

Proper valves

the

engines

linked

actuator.

oxidizer

instance,

small

mechanically

In high-thrust

on the

detail

control

For

Ideally, the basic theories

always be required in the process

Open-Loop

interlocks

reliability.

by a single

stage

The

is used

should permit design without

experiment or development

system,

condi-

as

accomplished and past experience

open-loop

for variable

sequencing

Mechanical high

propellant

With

it is

parameters,

fluid (electric,hydraulic, or

pneumatic), and of the operating mechanism

preset

of operating

accomplished

Interlock supply or working

new

set

control

However,

system, can be analyzed.

is an important firststep in control systems

system

Open-loop

operation.

Accurate

the functions and

selection of the best-suited method

design.

data.

of simplicity.

to a specific

is usually The

systems

forms of closed-

loop control. For these, mathematical

proposed

test

advantage

limited

dynamics,

from

control is confined

are designed

at a fixed, steady-state level over a narrow range of environmental

is determined

on-

of flow

of conditions. simply

Control

Closed-loop or feedback

III.

control is also called automatic

control. This

accurately sensed by

system usually in-

cludes sensing means, computing means to detect errors, and control means to correct them. An feedback

is compared

with a

fixed or variable reference by a computer, which 263

_mm

264

DESIGN

then generates

signals

to correct

I=

OF LIQUID

PROPELLANT

For rocket engine application, closed-loop control systems usually employ one or a combination of the following modes of operation: 1. Simple "on" and "off" type.-(Example: pressure switch/valve combination for tank pressure control.) 2. Proportional type.-Employs a continuous control signal which is proportional to the error. (Example: transducer output for chamber pressure control.) 3. Derivative type.-Employs a continuous control signal which is a function of the error and its time derivative(s) (rate of change). This is principally used when systems stability is critical. (Example: thrust vector control system with phase lead.) 4. Integral type.-Employs a continuous signal which is proportional to the cumulative integral of one or more errors. (Example: two flowmeter outputs for mixture-ratio control.) Closed-loop or feedback control systems are essentially dynamic systems. Their design characteristics may be analyzed according to the basic laws of physics. Figure 7-1 shows a typical example. Its function is to maintain the variable

Pc equal

_

ROCKET

ENGINES

for any devia-

tions. The main system thus does not require precise calibration for a specific set of conditions. Unlike open-loop control, closed-loop control depends on sensing absence or presence of an error to maintain a desired condition or to bring about a correction. In general, the objective of closed-loop control is to minimize errors during operation and reduce system sensitivity to environmental changes and changes in component characteristics. It is applied to areas such as engine-thrust control and/or throttling, propellant mixture-ratio control, and thrust-vector control.

controlled

mw

to the desired

value

Pr, by manipulating the variable wg. Maintaining Pc equal to Pr is assumed to maintain the indirectly controlled quantity F. In a typical turbopump fed engine control system, Pc would be the combustion chamber pressure, maintained equal to a fixed reference pressure Pr by means of a valve controlling the gas generator propellant flow wg. F then would be engine thrust, which is indirectly maintained at a desired value. In this control system which consists of a sensor (chamber pressure transducer), a computer

Figure

7-1.-Schematic control

of a typical system.

closed-loop

(electric summing junction and amplifier), and a controller (gas generator flow control valve), the command reference input r is compared with the sensor feedback b. The controller then manipulates • g in response to an error signal e from the computer. Ideally, r should be in linear proportion to Pr and b to Pc, save for constants required to convert one physical quantity into the other. However, this ideal condition is difficult to attain because of the dynamic characteristics of the pressure transducers. These characteristics are influenced by physical properties such as mass inertia, fluid compressibility and viscosity, and frictional resistance. Instead of r being directly proportional to Pr, the two parameters are actually related through a differential equation which represents the dynamic behavior of the elements involved. The same is true for the feedback b and the controlled variable Pc. It is also applicable to other systems components. Hence, the analysis of a closed-loop control system usually involves the solution of sets of often complicated differential equations. Refer again to figure 7-1, where Pc is the controlled variable, _¢g the manipulated variable, e the error signal, b the feedback, r the reference input, and Pr the desired value. A, B, C, and D symbolically represent the dynamic relation between input and output of the respective components. The following terms representative of the differential equations for this closed-loop control system can be written: r =Apt

e =r- b

Pc =C_vg (7-1)

b = Dpc

¢¢g = Be

The solution of these equations in combination with a systematic experimental program will suffice to analyze the dynamic performance of the system. The continuous corrective action of a closedloop control

system

may promote

dangerously

DESIGN OF CONTROLS AND VALVES

unstable

operation

components significant

response one

is

taining

a variable

is

no longer

for control

ity

difficult

often

The

high

ing

high

degrees

of system

results

possible

Higher

However,

leads

to obtain

satisfactory

dur-

various

such

(time

a high

tion on compensation will nection with thrust-vector

as

derivatives),

gain

stability.

through

control

it

system

Additional

informa-

be presented control.

in con-

safe as

The

engine well

cutoff

shutoff

and

As

sequence

is

a rule,

off in the

main

damaging

in smooth

Engine

(gas

generator,

propellant

rapid

LIQUID

ENGINE Most of the

basic

found A-4

CONTROL

engine

following

PROPELLANT

spikes

paragraphs.

in chapter

Ill,

propulsion

for the

systems

several

or all

summarized

Typical

in the

applications A-I,

(figs.

results

governing engine

been discussed

in section 2.i.

is the result of a malfunction, will be supplied

require

systems

and

pre-

The signal for engine in-flightcutoff, unless it

SYSTEMS

systems

control

ROCKET

cut-

This

termination.

Important consideratious BASIC

valve-closing a fuel-rich

Main Stage Duration Control

duration have 7.2

in

(purges,

chamber.

thrust

of

and,

securing

temperature

and

power;

to provide

combustion

opera-

consists

chamber the

cutoff

systems

postfiring

adjusted

noris

usually

power

of main

flushes).

repeatable

reliable

sequence

firings,

during

in an emergency,

to enhance

of test

vents

shutdown,

as

of subsystems

etc.); case

Control

for minimum and

shutoff

gain.

in overshoot

and

desirable

tion.

accu-

high

promoting

of compensation,

phase

The

Cutoff

operation

impulse,

for stabil-

i.e.,

instability.

means

"anticipatory"

with

to combine.

thus

in main-

in.

System

Rapid mal

Instead,

and

amplification;

action,

appropriate is

accuracy

amplification

corrective

set

and

control

value.

may

Engine

or gain

effective

desired

oscillations

requires

high

An unstable

at its

requirements are

elements

having

lags.

that

divergent

racy

control

employed

system large

when

are

265

A-2,

2-10,

are

A-3, 3-3,

and

3-6,

and

by the vehicle and fed directly into the cutoff control system discussed graph.

optimum

utiliza-

tion of the propellants is desired, a tank lowlevel sensor is often employed. where

3-9).

in a preceding para-

For lower stages, where

In final stages,

precise cutoff velocity is essential, an

integrating aceelerometer

or equivalent device

will signal cutoff. Engine

System

The trol

Start

prime

objective

is to bring

start

signal

the

may

(purging,

chilldown);

if required

consist

(start

introduction

and

required

the

generator

and

application

ignition

sequence

This

2-11, system

3-8,

start

of start

energy,

instability,

and

sors,

propellants

Secondary

in the

sequences

of engine

interlocks

control

maintained the

through

functional start

step

transient.

and and

and

by propellant cooling

3-11

cutoff

methods.

present sequences.

typical

fail-safe

of engine

tion

The

system

opening sequence is set to an oxidizer-lead or a fuel-lead start. dictated

ous,

type

and

system

systems

to shutdown

designed

during

it may

be desirable

gency

power

source,

will

Figures

latching, for continued and electrical interlock

engine

extensively

used

the

reliability

in the

of the

control

safety

an interrupcause

to an emerby

operation.) devices

system

control

the mis-

shutdown

mec_Janical Mechanical

phases

engine

certain

to switch

or prevent

all

most

(For

condi-

nonhazard-

so that

supply

safely.

sions

frequently

or unsafe

automatic,

shutdown

power

combussen--

are

In addition,

are

of electrical

trips,

undesired prompt,

operation.

as

combustion

overtemperature

overspeed

to prevent

such

such

for detecting

generator

by effecting

engine-

devices,

monitors gas

employed

Controls

monitoring

or turbopump

tions as

Safety

stability

subsystems

each

ignition 3-5,

Special tion

A reliable

is

is usually

chamber

A typical

System

preconditioning spinner);

of the

Engine

system.

during

propellant-valve effect either

from

for certain

by monitoring

operation

turbine

con-

safely

operation.

chamber.

may be start

system

of systems

tanks,

combustion gas

of a start-sequence

engine

to main-stage

sequence

main

Control

are

to assure

systems.

266

DESIGN OF LIQUID

Propellant

Tank Pressurization

Various tems

propellant

have

been

tank

design requirements tems must consider(1)

Means sure

(2)

within

an allowable

all

phases

of vehicle

(3)

operation,

transients;

periods

between safety

control

and

to prevent of the

with

such

as

pres-

range and

engine

steady-

tank

are

closed-loop

Engine

System

or dynamic

vehicle

coasting

other

as

pressure

overpressurizapropellant subsystem

tanks. con-

propellant-utilization-

and

thrust-control

Figure

7-2.-Control

for desired

and

ori[ice

for each

some

are

parameter

pneumatic stricting

sizes

switches,

are

checkout

ance

and

includes

of orifices.

of these and

cussed

locations

sizing

placed

the

and of timing

switches,

correct

values

during

engine

firings.

Of the

orifices,

Specific ratio 7.3

lines Others

lines

mixture

o[ a typical

setting

position

in propellant

in sections

para-

calibration

characteristics

The

calibration.

devices.

preceding and

verified

or hydraulic and

in the

adjustment

operating

This pressure

the

for thrust

systems.

described

engine

control

type.

Calibration

proper

performance.

pressurization

Control

require

bration such

of the

systems

graphs

devices,

start

restarts.

rupture

Compatibility trols,

sys-

tank

including

devices

valves and

required

mainstage;

throttle

tion

of these

level

Effective

The

propellant

systems

The

engine

relief

control

the

systems

sysV.

for the

ROCKET ENGINES

Most

in chapter

to maintain

during

state

Control

pressurization

discussed

PROPELLANT

and

engine

as

are

used and

in re-

applications

control 7.4;

for perform-

timing

orifice

cali-

will orifice

system.

be disdesign

DESIGN OF CONTROLS AND VALVES

elements

will

be presented

Following

sizing,

identified,

by stamping

actual Control sizes

sizes and

To

verify

means

are

firing

Checkout

its

components, system.

without

and

Provisions

of all

for verifying

voltage

monitors; position flowmeters;

checkout

to tile

ground-support

The

rocket

been

explained

specified

of the

include firings.

It is possible,

orifice

of the

fed

level

in section

2.1.

and

Main

(3)

Variation

for

tems,

engine to

often

of a has

usually

The trate

of the

engine,

various

propellant

without

resort

to regu-

of turbopump

regulation

of gas

rate

or hot

gas

method). rate

tank

pressures

(in

the

systems). (clustered)

reduction

or more

example

chosen

a closed-loop

engine can

engines

be

sys-

effected

of the

power

variation.

control

system

which

relies

Here,

the

thrust Figure

on

main

7.1

system

is

control

sub-

for our

the

flow system

for

turbine' thrust

A-4 stage

propellant control

to illustypical

through

7-3 shows

proposed

closed-loop

the principle main oxidizer

engine,

variation. operates

on

of variable fluid resistances in the and fuel feed lines to achieve

flow-rate by

parameter

most

comparing

the

reference

in section

control

effecting

is determined

simple

thrust

of one

propellant

in the

control

systems.

"_+3%."

calibrations

pro-

pressure,

by

case flow

flow

in multiple

level"

with

the

of pressure-fed

stepwise

"fixed band

the

basic

of Pc

through

of main

Additionally,

with

this

(in

propellant

case

for instance, thrust

rate

(2)

a tolerance; modern

power systems),

(preferred

or vacuum)

Two

be accomplished

flow

flow

CONTROL

It is

reduction can

as

level

is for a planned

of chamber

propellant

equipment

thrust

require range.

during

flight.

generator

This

need

or "throttling,"

reduction

Turbine

a system

LEVEL

(sea

to guarantee

subsystems

(1)

instrumentation.

THRUST

the

will

a wider

possible:

by shutoff

test

cases,

of them

func-

the

predict-

of-

for start

equipment,

missions over

Continuous

Each

system

as

to cham-

is

Pc (b)

signals

engine

such

same

at sea

of altitude

of propelled

are

is

The

of thrust

of thrust,

portion

ratio

relationship

control

(a._) Stepwise

devices.

vehicle

must

engine

in such

last

mixture pressure. starting

vehicle

Usually,

as

regulators,

system

with

engines,

entire

is

accuracy.

thrust

reduction

effect

at altitude

for systems

control

control

of static

significance

liquid

proper

high

and

pressure

their

since

a function

Occasionally,

firing

of all

ratio

checkout

additional

ENGINE

and

the as

in-flight

transducers.

pressure

operating sequence, and cutoff.

7.3

with

continuous

such

to simulate

"cold"

requires

able

propellant

the

subsystems,

mixture

Provisions

control

engine system

true

because pressure

equip-

plug and

range

valves, and

level, ber

such

pressure

operating and

thrust

permit

of the control

of the

for valves;

and

and

control

In addition

is essentially

operation

for verifying

devices

(4)

and

proper

spark

regulation, engine

pre-

at sea system.

chamber

(vacuum),

of chamber

pickups,

closed,

signals

Provisions tion

and

open,

actually

a function

cedures

instrumentation

dc bus

are

At altitude

for a given

system.

Provisions

regulators

solely

checks

checks

Thrust

a

starting

of a multistage

thrust

the

leak

stages

are

require

and repeatability, vehicles

or in final

to thrust

"controllers"

which

of precision

level,

identical

or

systems

in single-stage

regulators.

engine

control

to conduct

7-2.

as

of calibration firings.

regulators

control

critical firing

degree

such

Controls of the

checkout

higher

in figure

ground-support

electrical-continuity

(3)

their

simulation

its

actually

ment (GSE), an engine should include-

(2)

in vehicle

suitable

suitable

engine

employed

shown

permit

engine

Utilizing

(l)

and

readiness

These

of the

"thrust"

for postassembly

checkouts.

lators, and with a minimum However,

and Test

subsystems,

required

operation

are

operational

and

7.10.

be properly

in the engine logbook. orifice locations and

system

Systems

system

must or banding,

recorded calibration

of a typical

Engine

in section

orifices

267

modulation. sensing

indicative feedback

pressure

input

Engine

chamber of thrust b with r.

the

reaction

pressure, level,

the and

by

command

Any resultant

error

e,

268

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

7.4

C_lu_m _¢_E

,

_

o

_mu

,NPUT

[_

Figure

7-3.-Main-stage thrust loop Ior the ,4-4 stage

C_EC_m_

The significance of propellant mixture ratio and its control have been discussed in section 2.1. The principal reasons for mixture-ratio control are recalled:

control Open-Loop

following amplification and compensation as required, is used to drive the thrust throttle control actuator of the main propellant control valves in a direction which reduces the error. Ideally, the system operates over the entire thrust throttle range with minimal disturbances to other critical engine parameters; in particular, the propellant mixture ratio. In practice, these disturbances are not entirely avoidable, but can be minimized by maintaining a given resistance ratio between the two main propellant control valves throughout the control range. A most reliable method toward this objective would be mechanical coupling of the two propellant valves (fig. 7-4). Orifices, propellant valves, and servovalves required for thrust control will be described in subsequent chapters.

ENGINE

FROM

ST_.RT,

THROTTLE

CUTOFF

CONTROL

{I TPiRUST

_II£TL_TO_

7

FI_:_I

OXIDIZER TANK

._,

N OXIDIZER T _ CO_ ROL LVE

TO THRUST

NO.

I

CHAMBER

/

_

FUEL _'ANK

/

i

MECHAN ICdI, L COOPLING

TO THRUST

NO

Z

CH*_*M_R

TO THCtUST

NO.

I

CHC_MBER

TO THRUST

Figure 7-4.-Schematic of the propellant system for A-4 stage engine start, throttle and mixture ratio control.

AND CONTROL

Optimum engine performance (important) Complete propellant utilization; i.e., minimum residuals (most important) Both goals are closely interrelated and essentially inseparable.

i,,_._,_

throttle engine.

T

PROPELLANT-MIXTURE-RATIO PROPELLANT-UTILIZATION

NO

Z

CHAMBER

control cutoff,

Mixture

Ratio

Control

The simplest form of engine mixture ratio control is obtained by the installation of properly sized calibration orifices in the main propellant lines. Acceleration effects during flight are usually accurately predictable as a function of trajectory and flight time. Thus, simple averaging of flight mixture ratio and selection of the corresponding orifice size reduces mixture ratio deviations over the duration of flight to a level acceptable for optimum total propellant utilization in many missions. Open-loop mixture-ratio control can often be further refined by the following procedures: 1. Weighing ot the propellants loaded; i.e., accurate determination oI the tanked propellant mixture ratio.-The vehicle to be launched rests on load cells, thus permitting weighing of the propellants actually loaded. In mixed systems, the noncryogenic component is loaded and weighed first. The cryogenic component follows and is subsequently maintained at level through a topping line. The mass of both propellants is determined from on-the-spot temperature and ambient pressure readings while the tanking procedure is progressing. 2. Use o[ adjustable, rather than [ixed, ori[ices in one or both propellant lines.-As close to vehicle takeoff as possible, and as a function of tanked weight and temperature readings, a hand or remotely ground-controlled prestart-oriflee adjustment is made. This method is usually confined to noncryogenic fluids. For systems where engine operation closely follows that obtained during final calibration, remarkable accuracy of targeted mixture ratio and thus propellant utilization can be obtained

DESIGN OF CONTROLSAND VALVES

with the open-loop method, closed-loop system (single In certain applications,

approaching that of a stages; first stages). however, the varia-

tion of mixture ratio as a function of increasing acceleration may exceed tolerable limits. Acceleration in most vehicle tank arrangements affects predominantly the propellant in the forward tank. Because of the long supply line, acceleration continues to act upon a relatively large fluid column, even near the end of powered flight (tank depletion). By comparison, the effect on the fluid in the rear tank is often nearly completely offset by the simultaneous decrease in fluid head (short liquid column). To offset excessive acceleration effects on the fluid from the forward tank and thus on mixture ratio, head-suppression valves are sometimes used at the pump inlet of turbopump fed systems. Here, pump inlet pressure increase is sensed as a function of acceleration. Corresponding signals are fed through a logic device to the head-suppression valve which will gradually close, thus acting as a throttling device. This valve also protects the pump structurally.

Closed-Loop

Mixture

Ratio

Control

In certain cases, such as in last stages, or in missions requiring engine restart following extensive cruising periods involving propellant boiloff, a closed-loop system may be required. In figure 7-5 we see the A-4 stage engine mixture ratio control loop which operates on the basis of continuous propellant mass flow sensing. Both fuel and oxidizer mass flow rates are monitored and integrated to establish the ratio of either the propellants consumed or the propellants remaining. The mixture ratio feedback, (MR)b, is then compared with a command reference mixture ratio input, (MR)r, in the propellant utilization control computer. The resulting error signal, (MR)e, is fed to the mixture ratio control oxidizer valve

_9

oxidizer flow in a similar manner. For instance, a propellant-utilization servo control valve, which regulates the pneumatic pressure to the main oxidizer valve actuator, may control the oxidizer flow by adjusting the angular position of the oxidizer valve gate during engine mainstage operation. In certain applications it may be desirable to integrate the propellant flow rates and to compare the masses consumed to one another and to those tanked for optimum propellant utilization. It is readily seen that control systems, based on propellant flow-rate measurements, are a refinement of open-loop systems using fixed orifices. They are basically still mixture-ratio controls and thus merely "assume," but do not measure directly, the amount of propellants actually remaining in the tanks and their unbalance. To accomplish this function, usually referred to as "propellant utilization" (PU), additional control elements must be employed in the form of vehicle tank-level sensors. Numerous principles are known: point sensing, sonar, acoustic, radiation sensing, differential pressure, and capacitance probes. Figure 7-6 presents the propellant utilization control system for the A-4 stage propulsion system. The residual propellant quantities in the main tanks are continuously monitored, summed, and compared with a PU control reference in the propellant utilization control computer. Any error detected is used to modify the command reference mixture ratio input, (MR)r, to tl_e mixture-ratio control computer. This method isolates the mixture ratio control from the propellant utilization control, and thus prevents interaction between them. The bandwidth of the

ilXTU_E

mXTU_f

_*TIO

_A_

vernier position actuator, which forms a link in the mechanical coupling between the two main propellant control valves, as shown in figure 7-4. The oxidizer flow rate is thus modified to eliminate the error. In high-thrust turbopump-fed engine systems such as the A-2 stage engine, where the propellant valves are independently actuated, the system propellant mixture ratio control can be accomplished by varying the main

[CEC_e,( _UV.C'nON _'

Figure

eLecrllc

SUVltNG lm_t.,1,1Ei

7-5.-Propellant mixture ratio loop for the A-4 stage engine.

control

270

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

°°

]

ELECT't;

Figure

7-6.-Propellant [or the A-4 stage

utilization propulsion

AMPL=FI(R

control system system.

propellant utilization control system is made narrow as compared to that of the mixture-ratio control system, because propellant residual errors may be expected to develop slowly; i.e., initial tanking errors can be corrected over the entire duration of engine operation. The sensors used in the vehicle tanks may serve additional purposes. In combination with suitable ground equipment, they may permit an automatically controlled loading, high-level limiting and topping procedure. In static firings and flight, they may serve as redundant low-level sensors to initiate engine cutoff. For such a complete system, the term "propellant management system" has come into increased usage. Apart from throttle valves placed in the main propellant lines, bypass lines have been successfully applied to vary mixture ratio. Here, a line is tapped off the pump outlet and ducted back to the pump inlet. A servo valve, possibly supported by an orifice, can be varied so that the bypass flow is adjusted from no flow to full bypass flow. The implementation of closed-loop propellantutilization control through mixture-ratio control is a major vehicle-to-engine interface area. The requirements or criteria will usually be established by the vehicle builder and/or user. Close coordination between engine and vehicle designer is essential. A closed-loop mixture-ratioand propellantutilization-control system may not only be used for accurate maintenance of a fixed mixture ratio but it also has the potential for programed mixture ratio control (PMR). Here, the mixture ratio is varied during flight, either continuously or in steps. It must be kept in mind that the average mixture ratio still must be equal to the tanked

mixture ratio to assure simultaneous propellant depletion. However, by programing a mixture ratio in favor of the heavier component during the early portion of flight, and then switching it in favor of the lighter one, the accelerated vehicle mass is reduced faster. Also, mixture ratio may be programed to provide a higher thrust level during the steeper portion of a trajectory. This provides a better thrust-to-weight ratio in the presence of gravitation, with attendant velocity increase benefits. These methods, possibly in combination, may substantially increase stage payload capacity, since the effects of mixture ratio on performance (Is) are usually small within a reasonable range (see table 7-1). Optimization can readily be made with the aid of an electronic computer program. In a number of applications, programed mixture ratio control without PU control, i.e., open loop mixture ratio control with PMR, may give best results, simultaneously reducing complexity. Valves suitable for mixture ratio control will be discussed in section 7.8.

TABLE

7-1

Mixture i ratio, Thrust O/F 3hange, percent

+10 -10

General

Considerations

Design

The precision ratio is obtained

+11 -11

Flow rates NPSH

Is

+12 -12

-1.3 +I.3

OxiFuel dizer +4 -4

+14 -14

with which a desired mixture or maintained is affected con-

siderably in open-loop systems, and to some extent in closed-loop systems, by the following: (1) Instrumentation accuracies (in particular, flow and tank-level metering) (2) Machining tolerances of orifices (3) Operating tolerances of regulators (4) Temperature influences on orifices and regulators (5) Density tolerances of the propellants, as a function of temperature and of purity (composition according to specifications; contamination and dilution)

DESIGN OF CONTROLSAND VALVES

(6) (7) (8) (9)

(10)

Acceleration effects during flight Propellant tank pressure deviations Turbopump speed deviations Differences between fuel and oxidizer pump characteristics as a function of speed Line resistance changes as a function of temperature and for miscellaneous mechanical reasons

(11) Temperature effects in rotating machinery In the following we will discuss important steps toward maintenance of high quality, and toward further improvement in the listed areas, for highest accuracy of mixture-ratio (and propellant-utilization) control. First, continued improvement of propellant flow-metering devices is imperative. Here, turbine-type flowmeters have achieved a high degree of accuracy (conformance with truth) and precision (repeatability). The accurate calibration of these meters to most reliable standards engine inlet pressures

is important. Since also affect the mixture

ratio, pressure measurements of the highest reliability are equally necessary. Wherever possible, the rocket engine design should include vital metering and measuring elements from the outset. Dynamic sensing devices, in particular flow meters, are drastically influenced by their installation configuration. If these end organs, following accurate calibration, remain with the engine through its entire life cycle, including flight, a maximum degree of accuracy is obtained. The design and machining of all calibration orifices should closely follow accepted standards (see section 7.10). Selection of suitable materials to eliminate or at least to reduce to a minimum, temperature influences and corrosion, is important. The design of orifice holders must prevent the possibility of incorrect (upside down) installation and of distortion of the orifices. Regulators, if any are used, must be designed for highest accuracy and precision with particular consideration of the medium to be controlled. More detail will be presented in section 7.12. The purity and composition of the better known propellants are regulated by official government specifications. The designer can expect that approved sources will deliver the propellants in conformance with these. However, subsequent contamination, dilution or alteration is

271

always a possibility and must be prevented by proper design and handling procedures. Many of these, such as cleaning procedures, will be called out in the shop drawings. Furthermore, the design, where applicable, will have to include filters, check valves, and suitable line routing in order to prevent contamination and/or contact with incompatible materials. Note that some propellants may change their properties merely as a function of time, such as hydrogenperoxide, which loses its concentration due to (very slow) decomposition (with attendant gas development), even if absolute cleanliness has been maintained. This affects design conditions in addition to contamination considerations since proper venting devices must be provided. The latter, in turn, have to be designed in such a way that no contaminants, including moisture, can enter the propellant system. Since mass flow rates delivered by pumps and/or regulated by orifices will be a function of the fluid densities, mixture ratio may be affected accordingly. The densities, in turn, aside from conformance with specifications, will be affected by temperature (noncryogenic fluids) or ambient pressure; i.e., boiling point (cryogenic fluids). To overcome these effects, it may be necessary to temperature-condition the propellants. This may be done by heating or cooling. Or, it may be accomplished by suitable storage, such as shielding against solar radiation. For cryogenic propellants, it is usually sufficient to keep the containers vented to atmosphere until immediately prior to use, since the possible changes of atmospheric pressure at a given altitude can only introduce relatively minor temperature changes. The designer, through a suitable operating sequence (engine schematic) and through provision of vent valves, recirculators, heaters, and other components, can minimize temperature effects. The actuation

of mixture

ratio

control

devices

affects the nominal engine performance parameters. Depending on the type of engine, in particular its turbopump characteristics, these effects may be significant. In an actual case, the effects shown in table 7-1 were observed. It is clear that the vehicle thrust structure must be capable of absorbing the higher thrust loads. Also, the vehicle tanks and their operating pressures must be capable of meeting the

272

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

NPSH requirements for extreme mixture ratio excursions. Furthermore, chamber cooling may be affected. During sea-level testing, nozzles with high expansion area ratios may experience jet separation at the lower thrust levels (low Pc), resulting in vibration, destructive to engine as well as vehicle structure. Since vehicles are tanked for their nominal mixture ratio, and since engines are calibrated to this ratio, mixture-ratio valve excursions should be small for vehicles which are expended within a few minutes after takeoff. For stages, with long cruising periods prior to operation or reignition, and which use one or two cryogenic propellants, boiloff may have altered the ratio of the propellants in the tanks to such a degree that the PU system may be called upon to operate at or near its maximum excursion. It is, therefore, vital that the engine designer appraise the vehicle builder of all performance variations as a result of mixture-ratio adjustments, beyond the standard tolerances of the nominal performance values. Also, engine turbopumps must be capable of operating for extended periods with the valve in either extreme position. A propellant utilization system is a complex system. If required, it must be of the highest quality. Otherwise, it will do more harm than good. Only closest cooperation between vehicle and engine designer will assure optimum quality. Areas of particular significance to teamwork are: Selection of the mixture ratiocontrol method.For instance, should the PU system be active during the entire flight duration, or only for the last, say, 30 percent. (Both methods have been successfully used.) Selection of the mixture ratio control valve specilications.-Should it be a variable orifice, or a bypass valve? What should be the permissible pressure drops, required response rates, and accuracies? In case of sensor failure, should the valve return to the neutral position or remain in its last working position? (Self-locking.) Selection of the sensors.-Should it be one of several available continously reading types, such as capacitance gages or differential pressure (tank top to bottom) gages? Or should point sensors be employed, such as hot wires (change of heat loss as a function of being immersed in fluid or exposed); switches triggered magnetically by floats; voltage pips induced in station-

ary coils by a passing magnet, or others? (The engine designer will be involved in this selection only if the generated signals affect engine components.) Selection of the best-suimd electronic control system.-This will be largely influenced by sensors and control-valve selections. A propellant utilization system is not a malfunction prevention system. It does not add to vehicle reliability, possibly subtracts from it. Rather, it is a system required to live with a marginal preliminary vehicle design. It is a safe assumption, however, that the first flights of a new vehicle will not be for its ultimate mission. PU, therefore, will not be a vital necessity for these flights. Thus, enough time is available to thoroughly investigate, analyze, select, and develop the PU system. This time should be utilized. Both engine and vehicle builder have facilities and test programs to permit mutual exposure of their selected systems to flight and simulated-flight environment.

7.5 THRUST-VECTOR

CONTROL

To steer a vehicle over its trajectory, thrustvector control is applied. The following methods have found application: (1) Gimbaled thrust chamber or engine assembly (widely used) (2) Jet vanes (obsolescent) (3) Jetevator (4) Gimbaled thrust chamber nozzle (rare with liquid propellants) (5) Secondary injection (into the thrust chamber) (6) Auxiliary jets The first method is used most frequently, due to its inherent reliability and performance. The first four systems require actuators which may be operated by hydraulic, pneumatic, or electric means. The remaining systems are controlled by flow regulation.

Thrust

Vector

Control Systems

Using

Actuators

Figure 7-7 presents a simplified schematic for a thrust vector control system, employing hydraulic or pneumatic actuators. It may serve to explain the fundamentals of closed-loop thrust

DESIGNOF CONTROLSAND VALVES

toward the same end. Malfunction safety circuits are included to effect engine cutoff in the event of erratic operation. A typical schematic for a thrust vector control system using electromechanical actuators is shown in figure 7-8. Here, the actuator is powered by a continuously operating, constantspeed, 28 volt de motor, fitted with dry-powder metal bidirectional clutches. The control com-

V_HOJ C_OANCE CO_MANO _FE_ENC_

II

=° (_

ELeCTRiC #U_N

9,JMMPNG

_,.ECT RIC

Figure 7-7.-Typical control system actuators.

_PLIFJER

schematic o{ a thrust vector using hydraulic or pneumatic

vector control, even though the systems used in practice may differ significantly in detail. The actuators are controlled by commands, originating in the vehicle guidance system, which are a function of the vehicle's deviations from a prescribed path and of its response to corrective steering action. These signals are fed through an electronic thrust vector control logic to servo valves. In the system shown in figure 7-7, each servovalve modulates the fluid flow to its respective actuator assembly in response to an electrical error signal which is proportional to the difference between desired actuator position and its actual position. Feedback of the actual position is obtained through a transducer attached to the actuator. Additionally, tim actuating speed is sensed by a rate transducer and applied tO the control computer to stabilize the closedloop control through adequate damping. Instead of a rate transducer, electronic differentiation of the position

transducer

output

may be applied

r.lc

Figure 7-8.-Typical schematic tor control system using actuators.

273

_m,¢

lU_W,.G

for a thrust vecelectromechanical

puter consists of summing junctions and an amplifier as in the case of hydraulic actuators. The dc motor drives the actuator through the bidirectional clutches which are controlled by the error signal generated through comparing guidance command reference input with systems position feedback. To provide adequate systems damping, the actuating speed is sensed by a rate generator or through differentiation of the position signal. Apart from electrical feedback and compensation systems, mechanical feedback systems coupled with hydromechanical compensation "networks" are coming into increased usage. They are inherently simpler and thus offer higher reliability. Two basic types of hydromechanical compensating devices may be distinguished: piston-bypass devices and load-pressure-sensing devices. Piston-bypass devices utilize leakages past the actuator piston to introduce system damping and may make use of dynamic relationships to control time constants (a hole drilled through the piston is an example). Load-pressure-sensing devices, commonly called "pressure feedback" (PQ) valves or "derivative pressure feedback" (DPQ) valves, are widely used. Figure 7-9 shows a typical servovalve and actuator schematic with derivative pressure feedback (DPQ) and mechanical feedback. The only electrical signal required is the input to the "torque motor" (an electromagnetic actuator) resulting in deflection of the flapper of a differential valve, which drains to the sump. If the flapper is deflected, as indicated in figure 7-9 by the arrow, nozzle flow on side B decreases, with an attendant pressure rise. The reverse is true for side A. The resulting pressure differential forces the power-stage spool to the left, blocking the return line on side B, and opening it on side A. As a result, pressure Pb increases,

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

274

N"I

NOZZLE D4ERJVATIV_ FEEDBACK

IS, OL AT _0_

MOTOe

B

PlES$_RE LEAF

DERIV&TIVE PE'E$ SURE

•

TO_OU[

_TE_

-

SPRING

LO_D P_STON

DIRECTNDN FEEDBACK

P'_ST_ON

i_OT _',J

O$"

_ _WER

ST_

I, IEc:,I

_&

FEEDB, ACI( S_4mNG

_'OOL

RETURN

_IECH FEEDBACX

--

m[TUml

O_IFICE

ORF_.J[ _'N.TER

Figure

7-9.-DPQ

valve

wiLh actuator.

and Pa decreases, forcing the actuator piston to the left to apply the desired load force. Attached to the actuator piston rod is a tapered extension which acts upon the mechanical feedback linkage, including a roller and a spring. The mechanical feedback attaches to the torque motor. The pivot point of the valve flapper becomes the error torque summing junction. Note that the nozzle jets also have a feedback effect. The time derivative of the actuator motion, i.e., the hydromechanical compensation, is obtained through action of a derivative load pressure piston. This piston is affected by the same pressure differential that acts upon the actuator piston; i.e., by the load pressure. However, by inserting an isolation piston and permitting flow through an orifice bypassing the derivative pressure piston, the pressures affecting the latter can equalize. The degree of this effect is a function of the actuator pressure differential and its rate of change and of the bypass orifice size (shock absorber principle). As seen in figure 7-9, the derivative load pressure piston acts upon the valve flapper when displaced. Thus it provides the required time derivative of the actuator motion for compensation. As has been seen, it is possible to provide compensation in thrust-vector-control systems by either electrical or hydromechanical means, the latter now being often preferred for actuators. Conceivably, other control systems could be converted from electrical to hydromechanical networks. The analogies between the differential equations of the two network types often permit the use of existing electrical networks

and transfer functions by substituting the equivalent hydromechanical time constants. Table 7-2 may be found valuable by those who wish to familiarize themselves with some fundamentals in this field. Detail on the design of servovalves will be found in section 7.11. Demonstration

Example

Two basic types of electrical compensation networks exist: current output for voltage input, and voltage output for voltage input networks. Figure 7-10 shows a simple form of a current output for voltage input network. Find the analogous hydromechanical network. Solution The transfer work is

function

for the electrical

i I+RCS --:-V R

net-

Amp/volt

where i = electrical current (amps) V = voltage (volts) R : resistance (ohms) C : capacitance (farads) S : LaPlace transform operator (= j_o for sinusoidal forcing functions) From table 7-2, we obtain the equivalent hydromechanical parameters for i, V, R, and C. The new transfer function then is A

2

Q _1+ C_xPK)S AP

1 Cx

o V

1 Figure

G 7-10.-Current

output network.

for

voltage

input

DESIGN

TABLE

drop,

i, current

volts

coulombs/sec

component

' V = V A - VB

.......

AP,

Component

Q, flow,

dq _ =-_

Analogies

Describing

or quantity

pressure

275

VALVES

Hydromechanical

Describing equation

.........

AND

7-2.-Electrical-Hydromechanical

Electrical quantity or component V, voltage

OF: CONTROLS

drop,

psi

AP:

in3/sec

V _AP

PA - PB

dV Q = d--t V:

q = coulomb

Remarks

Analogy

equation

i_Q

volume

charge

_I_-------

V

Capacitor,

-------_

dv

....

[_----

i=C_-

Q

A p ----_

(Ap)2d(,_p)

= -g

Ap 2 ! Massless C _ --_ sumed

-_

piston

as-

--.._Q

farads

Ap : piston K :spring

area, in 2 constant, lb/in

V _=_-

Ap

Q_Cx_,A-'_-_

CSIIE3

_CxA

R _

P

1._

Parabolic linearized

CX

flow curve about

operation

point

Q ---...,lllb Resistor.

ohms

inS/sec --

Orifice,

V -----_

Piston

,, p

....

AP2 y(AP)dt Q:"M-

b

"Ap"

mass

not

negligible

t ------DInductor.

henries

Source:

Ap= piston M =piston

D. h. Engels, mechanisms,"

"A Method Proceedings

of Synthesizing of the IEEE,

Q area, mass

in 2

Electro-Mechanical PTGAC, October

The correct hydromechanical network, which is of the piston-bypass type, is shown in figure 7-11.

Compensation 1964. The

be

significance

seen

from

would Interfaces

With

Actuator

the

at the randomly tolerances,

Engine-to-Vehicle

Networks

be

for Hydraulic

of good

thrust

Servo-

alinement

fact that in an engine distributed

maximum

a trim deflection required

from

can

cluster, of these

of close

all engines

to 0.5 °

to offset

the

misalinement.

Systems For

larger

(looser)

trim deflection Engine For ance that

minimum and

the

alined point

engine engine

with in

tolerances cally.

Installation

all

and

demands

respect three are:

on

actuation thrust

the

planes. t0.25

the

vehicle

systems, vector

to

Alinement

inch

be

it

guid-

trim deflections

thrust

and

is required pre-

ered

properly

vehicle

attachment

Typical

specified

laterally,

if the

need

_*0.5 ° verti-

to

guidance

apply

flight

alinement

would

them

results

be

seem

to reduce

capability for

tolerances,

the

further increased.

the

full

in

appreciable

to

aline

only

Even

effective slightly,

duration

of

payload

re-

ductions. It

is

vector prior

customary to

to

the

upper

shipment•

the

engine

the

gimbal

face

of

Both

optical

and

thrust bearing dynamic

the

pow-

DESIGN OF LIQUID

276

PROPELLANT

ROCKET ENGINES

Q _C-_

/--VECHICLE

//,4i7///

_7

THRUST

MOUNT

CENTERLINE

/

._//

/

/-_..,_ cE.TE_ ,

-/- ,'.t-_.._q EAANCDH yIA_/

Pt:CAHE

)

F_

__

_

GIMBAL

"ARING

to T_IJ_

P_llel

\

\

\

[

Actu_

v_14m

• B * COITe_I_

L.el_gl.h :A

_tCCD

\.\\

AP

\

\

I

"

\

),4 Cx

-, \ "

_--,.,UST ,,c,o,

I,=, Figure

7-12.-Engine

/l\

alignment.

"o "_

x'x-

S

_L_ Figure

7-11.-Piston

methods

(load

cold

alinement

tion

of the

finding exit, dicular

actuators.

have

been

As

meeting

alone.

The

actuator index

can

be marked

gimbal had

(fig.

or as 7-12).

in a suitable

bearing

been

eye-to-eye

points,

points

properly

can

face.

alined

distance the

Lateral

mating

to the

means

engine

usually

vehicle.

produced,

it

of engine to the

launch-

of the

hydraulic-piston

electromechanical, types

other

(fig.

vehicle must ing

and

have

They

also

been

reserve.

end

As a rule,

tant

to note

that

dimensions

effected

by an

deflection") ducts,

flex

builder,

must two

actuators they

if the

maximum

individual

of absorbare

of a pair

is

gimbal

deflec-

It is impordeflection

angle

for instance, ("corner

approximately bearing,

required

permit

actuator,

maximum

and

an adequate

in all directions.

combined

lines,

with

by the

attachment

be capable

Together,

through

is 7 ° , the

procured

engine

The

to the at the

may be

encountered

engine.

attached

vehicle

dimensioned.

at either forces

are

to the

or by the

be properly the

and

7-13).

builder

engine

face

actuators

end,

of the

axes,

type

transportation

gimbal

at one

tion

at or on the

vehicle

the

verification

turbine-driven

for each

simply

vehicle

installation

into

of a new

Hydraulic-rotary,

points will

of the

line

manner

If the

of gim-

engines

be

shows

Loads are

Engine in

experience

by optical

alinement

the

attach

a few

manner,

specification as

two

after

vehicles

consists logbook

investigated.

confirmed

in lieu

7-12 engine

following

pneumatic

during

may be cells

type.

used

simply engine

site.

Actuators

plumb

may be

the

to specify

Actuators,

perpen-

A simple

load

first

or

nozzle

Subsequently,

in this

vertical

documented

line

center

a rule,

alined

permit

plane.

and

then

Figure

loca-

through

throat

connecting

side

the

alinement

optical

engine

of a prealined

is advisable

shop,

measurement

using

For

geometrical

in the

operation. this

methods

Tile

the

injector

of this firing,

dynamically bal

gimbal

to the

it, observing

ing

of nozzle their

of the

of attaching specifications.

hydromechanical

used.

vector

centers

to the

support

are

establishes

alining

attached

engine

cells)

thrust

the and

bypass network.

installation

10 °. and

possibly

Inlet

277

DESIGN OF CONTROLS AND VALVES

,_-_

ACTUATOR_ --

-- ..... SHORT

/-"

LONG STROKE ACTUATOR

where

filling

during

buildup

of the

duced,

they

these loads AOTUA.OR --\ ".CON,,OORAT,O

\

STROKE

_-j_

actuators

R ,\,.\\

can

Since

OlMBAL

CENTER_

I I

--I

[i

I I

entirely

1

situation

I

ator Figure

7-13.-Engine

actuator

installations.

and

in one

plane.

components

"take"

this

limited, bers)

affected,

deflection.

proper

be able

to

capability

gimbal

restriction

be provided

(circular

gimbal

of the

actuators

must

instead

must If their

plane

(stops

or snub-

of square).

Selection on the

gimbal

the

actuator

gine

thrust

design

forces force

required.

the

Inlet

be 25 percent

The

force

duct

Heat

of the

is determined

ture,

if it has

by

does

shield

reactions

friction

reaction

Correction

(if any)

for misalinements

Aerodynamic Vehicle

loading

It is

stressed

size

these

that forces

and lightest

actuators, tion

of this

toward

this

of stable,

and

goal.

such

as

though

the

loads

different

from

This

engine

design

of hydraulic

careful

The

system

dual-load

Recognican

must

do much

be capable

when

prelaunch

encountered

those

cold

checkout, here

may

occurring

during

situation

may

gimeven

be quite

engine pose

fir-

serious

problems. During loads

startup

in excess

gimbaling itself.

can This

sion

area

(for

engines

of the of those

be generated is

especially

nozzles designed

being

engine

deflection,

sation

network

engine,

brief

occurring by the true developed

for altitude

peak

side

during

and

deflection

per

low

mass system is

and

compen-

low

analysis

conducted

designer

will

stiff-

the larger

hydraulic

system.

is

high).

in low

actuator

effective size

However,

so is

the

and the

stiffness. and

a detailed by engine

which

effec-

in degrees

stroke,

jointly

has

high

requiring

is bulky Only

actuator

However,

results

reduces and

and

expressed

determine flight

engine

rate

dimensions.

members.

for the

degree

configuration

actuator

arrangement

structural

oscil-

per

and

typical

powerful

design

hydraulic

best

gains,

spring

(gain,

inch

gimbaled rate

stroke

is high,

a more

long-stroke

the

as with

oscillation of this

short-stroke

mass is

over,

motion A delay

continued

of compactness.

Resolution

spring

actuator

frequency

two

of high

advantage and

i.e.,

engine

requires

Moreextra

design and

vehicle

configuration

is

system.

normal

thrust

for high

shows

the

gimbaled

"soft,"

the

system.

feedback

7-13

tive

If struc-

parameters. The

The

to an

its

engine is

of actuator

installations. ness,

cross-

deflection.

(lb/in),

and

a function

actuators

pump,

design

response

during

should

for smallest

equipment.

well-damped

baled,

ing.

the

to a minimum

associated

need

is

Figure

effects

weight

and

lation

engine

to translate

engine

natural

develop.

between

guidance

The

If it

may

structure

overshoot

result.

in its

to minimize

rate

promptly the

actuator

by actuator,

thrust

a actu-

"crosstalk."

be able into

not

of the

other

builder

spring

react

for by

may

(if any)

acceleration

Inertia of gimbaled mass Miscellaneous minor effects

reduce

vehicle

subsequent line

bearing

not

called

motion the

formed

a low

are

instabilities

delay

loop

and

of

(notifi-

installation,

to as

must

without

control

en-

reactions service

the

case,

following:

Flexible Gimbal

In a typical

may

level.

considering

is based

ends

designer

coordination

actuator

motion

and

affects

is referred close

The

design

Rate

wherein

designer and vehicle talk is essential.

pattern

at both

actuator

control

Therefore, is

re-

for the

vehicle

exist

is excessive, other

Spring

regarding

This

or at least

points

and

may

Unless

builder).

engine free

time

ones.

considered

attach

Crosstalk

a longer

shorter

eliminated

be

of vehicle

takes

with

be

must

and

cation

nozzle

than

chamber expan-

at sea operation),

level

System

Hydraulic Until system actuators.

other is

means

probably Its

basic

are

available,

required elements

to power are:

a hydraulic the

engine

278

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

Hydraulic pump Reservoir (low pressure, or "sump") Accumulator (high pressure) Servovalve Actuators Feedback (electric or mechanical) Lines, check valves, filters, connectors, instrumentation If continuous hydraulic power is required prior to engine start, such as for recirculation of the hydraulic fluid or for gimbal tests, an electrically driven auxiliary pump is also provided. In most instances, the auxiliary pump will be operated until vehicle liftoff only, and can, therefore, be ground powered. For upper stages, the accumulator will then provide, for a limited time, the hydraulic power required during staging and turbopump buildup following its unlocking. Since some of these components will be part of the engine system, while others are stage mounted, an important vehicle/engine interface exists. Through an auxiliary drive shaft, the main hydraulic pump may be driven from the engine turbopump. It is connected to the other hydraulic equipment and to the actuator through high-pressure lines, several of which must be flexible. These other elements may be mounted on the vehicle at the expense of longer lines which also must cross the gimbal plane and must therefore be flexible. Or, they may be engine mounted. This, however, increases the engine gimbaled mass and may pose space and envelope problems. To compensate for misalinements and thermal expansion and contraction, a certain amount of flexibility must be provided for the lines even in this case. It is possible to connect an electric generator to the main turbopump, and drive electrically a stage-mounted hydraulic pump. Only electrical wires will then cross the gimbal plane, with the exception of the hydraulic lines to the actuators which always must be flexible. Another possible simplification is to combine servo valve and actuator into one single unit. Figure 7-14 shows a typical hydraulic engine actuation system. Figures 7-15 and 7-16 may serve to identify the major components of this system. From the above it becomes apparent that numerous hydraulic connections will have to be made when installing the engine into the vehicle. All of them must fit, and permit adequate flexure,

Figure

7-I4.-Engine

actuation (hydraulic).

1

system

1

schematic

...... PulIP

'--'=-'

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7"15.-Accumulator-reservoir

TRANSDUCEI

schematic.

PRESS.

SUPPLY

RETURN PUMP

SUCTION

EXTEND

_1

RETRACT 9.

FILEEI

16

BLEED

17

SERVO-ACTUATOR

II.

PREFILTRAIION

19

SAMPLING

20

VALVE

DIFFERENTIAL

21 . CYLINDER

26

Figure

7-16.-Servoactuator

SERVO

23

SNUBBER

24.

PISTON

2S.

LOCKING

26

MECHANICAL PRESS.

PRESS. BYPASS

22.

30.

VALVE VALVE INDICATOR VALVE

VALVE

BYPASS

VALVE

MECHANI_d4 LOCK

TRANSDUCER

31.

[_:.W.

32.

FEEDBACK

TRANSDUCER TRANSDUCER

schematic.

must be long enough and of the proper pressure rating, and, above all, must have a mating part on the vehicle.

DESIGN OF CONTROLS

Furthermore, consider system

perform.

and

ature

and

the

will

The

and

disconnect

are

at liftoff, method.

hydraulic

fluid

tioned

ground

electrically

is

of this

systems,

vital

heating

of the hydraulic supplied

heaters,

a favored

by means

and

which temperature-

Continuous

of the

the

cleanli-

temperature

Groin.

electric

temper-

Thus,

components

required.

com-

very

reliability

thermostat-controlled

conditioning

and hydraulic

extreme

engine

sensitive

is often

are

adequate

On cryogenic more

the

sensitive.

for maximum

system. system

fluid

to specify

requirements

know

narrow-tolerance

hydraulic

have

conditioning

must

in which

contamination

designer

of the

designer

environment

will

ponents

ness

the

the

279

AND VALVES

recirculation of the

driven

aforemen-

auxiliary

pump

another.

Secondary

Injection (C)

Thrust vector control through secondary injection of matter into the thrust chamber nozzle (SITVC)

has

motors.

It has

been

experimental for upper The

only

in liquid

where

it appears

stage

engines,

required

are

principal

applied limited,

application

systems, forces

successfully

found

methods

factors:

propulsion

mentation

promising

in which

the

than

7-17.-Secondary

amplification (K,).

(2)

Gas

injection,

(a)

Inert

(b)

Thrust

(c)

Gas

boosters.

K=ISs

injection

Liquid

Other

injection

(a)

Inert

(b)

Propellants such been

In a gimbaled located

With an located

required

side

The

at the

of preheated but the

side

injector side throat

of any based

type upon

¢Is

=secondary

¢¢p

=primary

Fs

=side

(7-15) fVs/fV p

end. force and

force,

is

rate,

rate,

lb/sec

lb/see

pounds axial primary thrust, increase, pounds

= undisturbed primary

Is s

= side

Is a

= secondary axial specific onds) = AF a/fV s

of fluid

and

the

the

K_ factor system

force.

If both

total tem

effect on

specific (seconds)

impulse

of a given

a propulsion

the

side

the the

factors

(sec-

may

the quanforce,

penalty required

are

secondary system

Fs/_'s

determines

determines

of these

impulse of the = Fp/fyp

impulse

to obtain

Is to obtain

pounds

(seconds)=

K factor

required

overa_A

of secondary

axial chamber

specific

Essentially, tity

two

performance

flow flow

Isp force

Evaluations

is

(7-1a)

fVs/_V p

Fp/_'p

Fp =undisturbed AFa = axial thrust

proven

force.

system

as

where

at the point of injection, resulting moment arm which decreases the

performance

aug-

7-17A)

using-

injection

chamber,

approximately

Performance

injection

7-17C),

SITVC system, the applied downstream of the nozzle

approximately in an increased

(fig.

7-17B)

investigated

thrust

thrust defined

lsa_AFa/ws_AFa/Fp tapoff

as

Fp/_Vp

=lsp (fig.

(fig.

are

K,

fluid

methods,

hydrogen, have uneconomical.

is

chamber

axial

Fs/fvs_Fs/Fp

lsp

gas

generator

and

factors

follows:

using-

stored

(K)

These

are(l)

systems.

injection

lateral

with

of secondary

INJECI'O'R

to solid

predominantly

especially

smaller

Figure

LIQUID

known, injection

on the side the sys-

be determined.

28O

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

The K factor determines the quantity of secondary injectant fluid required (for a known duty cycle); the maximum flow rate; the additional tankage, pressurization fluid, and secondary injection hardware weight; and the effect of the added inert weight on vehicle trajectory. The K 1 factor evaluates the Is penalty on the propulsion system. If K_ is equal to 1, the specific impulse of the secondary fluid is equal to that of the primary fluid and, therefore, the propulsion system suffers no specific impulse penalty due to the SITVC system. Both the amplification factor K and the thrust augmentation factor K1 are influenced by the secondary injection orientation. For each application, a tradeoff must be made between the two factors to determine the optimum injection orientation for maximum propulsion efficiency. Let the force of an external jet of comparable geometry at right angles to the primary nozzle be unity. Then amplification factors greater than 2 are obtained if secondary injection is made with the nozzles pointing upstream, rather than in a normal or downstream direction. Side forces for a given _Ps are further increased if injection is made through a series of holes arranged on a horizontal arc, rather than through a single orifice. Note that the manifolds required in this case may adversely affect response, however. Test experience suggests that overall pressure ratio and injector size appear to have little effect, while gas temperature does, optimum values being a function of propellant combination. For an oxygen/hydrogen tapoff system, the range between 3000 ° and 4000 ° F appears most favorable; however, as with turbines drives, material strength and cooling problems will dictate values substantially lower, say 1800 ° F. In a typical tapoff SITVC system, the gas flow rate may be 1.5 to 2.5 percent of the primary flow rate, the upper value indicating the situation of maximum force required between two injection stations (two jets operating). The tapoff system offers simplicity and good performance. However, with low-duty cycles, a continuous bleed may be necessary to maintain temperatures at the valves. The performance of a gas generator SITVC system is comparable to that of a tapoff system, probably slightly better. This is offset by higher complexity (valves, injectors, ignition, cooling).

Liquid injection systems (inert fluid or propellants) offer the simplest arrangement. This is offset by their low performance, K-factors being in the order of unity, at flow rates from 5 to 6 percent of the primary flow. However, in systems with low-duty cycles, they may still be very attractive. As a rule, four elements are required for a given system, equally spaced on the main chamber circumference, of which no more than two adjacent ones would be operating at a given time. The control of the required valves is accomplished through a logic and a servosystem analogous to that of a hydraulic gimbal actuator system.

7.6 DESIGN CONSIDERATIONS FOR FLOW CONTROL COMPONENTS

FLUID-

By theirvery nature,liquidpropellantrocket engines use many control elements for regulating and measuring of fluid flows, such as valves, pressure regulators, and flowmeters. Some of the design considerations governing these cqmponents are discussed below.

Basic

Flow

Characteristics

of an Ideal Fluid

Fluids, by definition, include both liquids and gases. A liquid is an incompressible fluid which is characterized by a tendency to retain a fixed density or volume; but not shape. A gas is a compressible fluid which has no tendency to either a definite shape or volume. Its density or volume will vary according to the basic gas laws (eqs. 1-9, 1-12, and 1-13). In general, the same fundamental laws of force, mass, and velocity apply to matter in all forms, and thus are also applicable to the flow of fluids. The analysis of fluid-flow controls may be simplified by initially assuming ideal conditions. For the calculation of physical dimensions and functional characteristics of specific control components, the results can then be modified by additional assumptions and empirical factors, which often are the result of extensive testing. A frictionless (zero viscosity), incompressible fluid which is nonturbulent and loses no mechanical energy as heat is referred to as an ideal fluid. For steady, ideal-fluid flow in a closed conduit, Bernoulli's energy equation applies:

DESIGN OF CONTROLS AND VALVES

144pl z[ +-p

to at least 10 times its diameter for repeatable

V2

V2"I

+

= z2 + __ 144 P2 +_-=2 p

2g

Assuming z t =z 2, and sions, we obtain

constant

rearranging

(%2)

the

expres-

results. For liquid flows, this flow-measuring method is fairly accurate, if frictionlosses are compensated

for by the velocity coefficient Cv.

For gaseous

flows, however,

pressure and tem-

perature have a significant influence on the den144 (Pl - P2)_ p

fluid

In conformance flow

sity of the

v22 - v,_

(7-3)

2g

with

the

continuity

law

fluid

and

must

be taken

into

account

for calculations.

of OtFFE_ENTIAL

_IAPH_

C,kA

/ PrLOT

# = Cvpv,A 144

1 _ Cvpv2A2 144

_¢ALVE

(7-4)

and

vl

A2

v2

A_

(7-5) _ENTURI

where

_ATE

z z, z 2 =elevations P,,

P2 = static

at sections pressures

tions v,, p

of the

fluid

flow

=venturi

rate

This

Figure

3

flow

ft/sec

velocity

configuration

2

the

coeffi-

of the

which

de-

is determined

tests.

conduit The

above

be used

flow

control

striction, is inserted provided

areas at sections

basic

systems. as

in the for reading

of the I and

fluid-flow

to measure

such

the

An accurately an orifice,

the

2, in 2

characteristics

or sense

conduit.

fluid-flow

nozzle, Pressure

static

pressure

flow

rate

sized

and

to control

in

re-

is

differential

between

venturi

throat

2).

springs,

butterflies, between

sure

The

area

design

factors

degree

A 2 and the fluid density p are known,

tions,

including

velocities v, and v 2, and the flow rate _¢ can be

spring

forces,

calculated with the aid of equations

accuracy.

preceded

by a straight length of pipe equivalent

by

means

discussed

of orifices

usually

obtained

in subsequent

relapreslinear.

permit functions. will

permit

in a reasonable Flow-bench of,

flow

regulators sections.

calibra-

for instance,

increase

of fluid and

the

exactly

dynamic

to further

control

charac-

venturi

analyses

adjustments serve

1) and diaphragms,

parts, not

accuracy.

Its

pressure

(sec.

of these thus

to the

dynamic

w and is

actuator

ports,

resulting

of control

The

inlet

other

(p_ -P2)

The gate

valve.

by the

rate

theoretical

of the restriction (sec, 2). If the flow areas A I,

venturi or orifice meter should be

and flow

calculations

a pilot

sensing

approximations empirical

a venturi flow.

pressure

of the

venturi

differential

and

(7-3), (7-4),

Because

tionship

good

the flow

by

controlled

of the

fluid

fluid

controlled

teristics

in which

across

by a fluid-powered

is

(sec.

system,

of a butterfly

working

in turn

However,

are

by means

The

diaphragm

or venturi,

p,

differential

used

of a typical

control

is positioned

taps

P2 at the inlet (sec. I) and at the minimum

and (7-5). The

closed-loop,

system.

schematic

pressure

diaphragm. by

the

is regulated

position

A 2 =cross-sectional

can

static

flow

is

fluid-flow

is sensed

and of the fluid-flow and

7-18

closed-loop

fluid-flow,

is a function

characteristics,

A_,

of a typical

at sections

32.2

of the

or orifice

cient. sign

lb/ft

constant,

= weight flow ib/sec

Cv

7-18.-Schematic

Figure

fluid-flow control

fluid,

=gravitational

IP

at sec-

2, fps

= density

g

fluid

2, psia

of the

1 and

2, ft

of the

1 and

v 2 =velocities

1 and

this and

pressure

will

be further

282

DESIGN

Sample

Calculation

OF

LIQUID

PROPELLANT

ROCKET

A

=area

#

=viscosity (viscosity

(7-1)

The followingdata are given fora horizontal venturimeter,measuring liquidoxygen flow: Venturiinletdiameter,dj = 6 in Venturithroatdiameter,d2 = 3 in Venturiflow velocitycoefficient, Cv = 0.92 Pressure differential between inletand throat (Pl - P2) = 22.5 psi Density of LOX, p=71.38 Determine flow rate w.

A2

2

into equation

V2

=_

(7-3):

Substitute

2g

2× 32'2 ×i44

71._s1-

this

1 × 2i'5-55.9

into equation

fps

turbulent flow velocity distribution is more uniform across the conduit than with laminar flow.

(7-4): 77

Flow rate

Cvpv2A 2 ¢¢=-144 = 180.2

Real

Fluid

Flows

0.92 x 71.38 × 55.9 ×-_-x 9 144

lb/sec

Involving

Pressure

Even in turbulent flow there is always a thin layer at the conduit wall, the boundary layer, which moves as a laminar flow. Experiments and theoretical considerations have shown that the Reynolds number, R e of a given fluid flow can be used as a criteria to indicate whether a flow is laminar or turbulent.

Drops

All real fluids possess the physical property of viscosity; i.e., they offer resistance to shear stresses. The viscosity of the fluid directly affects friction. The basic correlation is given by Newton's law of viscosity (see fig. 7-19):

mA u

F

/ /

I

/

I

= shear stress = F/A, lb/ft: = shear or friction force of the fluid tangent to the surface in consideration, lb

I

/

I

l

I

/ t

where r F

I

/ (7-6)

-Figure

ira,

/

/ gU r =-gt

poise)

the fluid moves in layers, or laminae, one layer gliding smoothly over an adjacent layer, with only a molecular interchange of momentum. The velocity of the fluid is greatest at the center of the conduit and decreases sharply to zero at the conduit wall. As the flow velocity is increased above the "critical" point, the flow becomes turbulent. In turbulent flow an irregular random motion of the fluid exists, in directions transverse to the direction of the main flow. The

144 (p_- P2)_ v22_ (,_v2) 2 p

2 = 14.84

When a fluid is forced to flow through a closed conduit, its flow is laminar or nonturbulent below certain "critical" velocities. In a laminar flow,

1

V 1 = ¼V 2

this

= [email protected]/ft2

=velocity of a fluid particle at the surface in consideration, ft/sec t =distance from the point where the velocity of a fluid particle is zero, to the surface in consideration, ft U/t =rate of angular deformation of the fluid

v2-A , \d,/ =-4

Substitute

ft 2

lb/ft-sec

U

lb/ft 3

(d2_

in consideration,

of the fluid, conversions:

- 4616.81b-see/in

(7-5):

VZ

of the surface

1 lb/ft-sec

Solution From equation

ENGINES

I _-veuoc_tv 7-I9.-Angular

OF FWtO PArTtcue_s zeao deformation

of a real

Iluid.

DESIGN OF CONTROLS AND VALVES

(R e : Dvp/tz, conduit,

where

ft;

D = equivalent

v=flow

velocity,

sity, lb/ft3; and/_=fluid For most calculations, flow

is

laminar

1200,

and

than

1200. Real

fluid

caused one

by rubbing

there

is

in pressure

in the

energy

is

heat

loss thus

may be extreme

be entirely

conduit

wall,

temperature

other

or isothermal

ambient

through

However,

flow

pipes

adiabatic

place

in nozzles,

valves

through

flow

is

orifices, which

by

to be

When

is

and

number

but

isothermal.

or tube

walls.

flows

may be

gram'

shown

at high

term,

pressure

drop

Ap (psi)

conduits

zontal

position

(7-7).

This

(4-32),

except

(ducts

can

of a fluid

or tubes)

be estimated

is essentially for the

the

flowing

in a hori-

by equation same

as

duct

control For

the

PV2 288 g

roughness"

to the

(e/D),

diameter.

components

friction

factor

modified

by

is a function

is

given

a measure

f obtained

of

values

in table

from

figure

7-3. shapes,

7-20

correction

Reynolds

of

engine

or for other

an empirical

dia-

projections e for rocket

are

of the

Moody

Average

passage

duct

dimensionless

roughness

projections

a curved-flow

of the

of turbulent

of the The

the

the Reynolds

factors

7-20.

roughness

upon

roughness

by means

surface

which L

found

relative

to be Ap:f

the friction

surface

flow

dimensions.

upon

of the

equation

(R e > 1200),

not only

in figure

"relative size

turbulent

Tile

and

moving

is

depends

also

to take

the

The

flow

factor

velocities.

in straight

the

friction

tubes,

fluid

(7-8)

f__6_4_4 Re

the

of liquids

assumed

number, and

The

Generally-at

short

the

can be arrived at by Poiseuille's equation for laminar flow

(constant-

is assumed

fluids, other

If the flow is laminar (R e < 1200), the friction

flow),

through

flow).

temperature-the

gases

sure at the outlet point. To calculate higher

drop

absorbed

extreme

is less than 10 percent of the fluid static pres-

factor is a function of the Reynolds

This

(adiabatic

dissipated

in the

with com-

the pressure drop Ap

Con-

energy.

entirely

in

itis recom-

that equation (7-7) be used

pressible fluids only where

against

flow.

Also, there will be a slight change

pressure drops of compressible methods should be used.

i.e.,

heat

case

mended

than

wall.

of the into

the friction factor. Consequently,

greater

particles

conduit

direction

ciably.

den-

friction

of energy;

converted

in one

or it may

the

a loss

produced

fluid,

fluid

against

less

numbers

involve

of the

and

sequently,

the

always

of the

lb/ft-sec.) that the

numbers

for Reynolds

flows

another

p=fluid

viscosity, it is assumed

for Reynolds

turbulent

diameter fps;

283

has

factor,

number

and

(7-7)

where L = length

of the

p = density

conduit,

of the

v = flow-velocity d = equivalent

TABLE

in

ness

fluid,

lb/ft

of the

fluid,

fps

of the

duct

diameter

3

Control or tube,

7-3.-Average Projections Component

Values for Rocket

of

Surface

Engine

Rough-

Fluid-Flow

Designs

in IRoughness

4 × duct

cross-sectional

area

Wetted I : friction

factor,

Equation flow tubes.

(7-7)

of any With

used

when

dled.

The

suitable density

great,

if the density

as

for laminar fluid

fluids

of compressible a function

pressure

drop

and

velocity

Surface

description

projection, _, ft

or turbulent

in ducts

or

it may

also

are

being fluids

of pressure; between will

Drawn tubing with very clean surface ........ Smooth machined and clean surface .......... Machined or commercial cold-rolled surface...

experimentally

restrictions

compressible

considerably fore,

is valid

shape)

perimeter

determined

incompressible

(any

two change

be

Rough machined surface .................... Smooth cast or forged surface ............... Commercial cast, forged and welded surface

..

0.000005 .00001 •00005 .0001 .0003 .0008

hanchanges there-

points appre-

is

1Moody, L. F., Friction Trans. ASME, Nov. 1944.

Factors

for Pipe

Flows,

284

DESIGN OF LIQUID

Figure

of the

design

creased can

configuration.

resistance

of a specific

be accounted or equivalent

which

is arrived

length

and

(L e + L), is then sents bends.

the

used

typical

Because

The

actual

flow.

resistance

in-

duct

sum

of this

passage

length,

(7-7) Figure

for the 7-21

characteristics

flow-control

diagram.

passage

of straight

in equation

of turbulent

ROCKET ENGINES

to it a ficti-

at empirically.

length

calculation

Le,

7-20.-Moody

the

flow

for by assigning

tious

equivalent

Sometimes

PROPELLANT

pre-

of 90 °

components

such

/ /

as

valves and fittings disturb the flow pattern, they produce an additional pressure drop in a duct or line

of tubing.

The

a flow-control sure

drop

component within

the pressure stream

loss

ducting

the

of pressure cQnsists

component

produced of the

itself,

as

preswell

as

drop in the upstream and downor tubing

in excess

of that

which

would normally occur if there were no component in the

line.

/

by

With

certain

exceptions,

the

fluid

Figure

7-21.-Typical

resistance o! 90 _ bends.

characteristics

DESIGN OF CONTROLS

flows are

through

rocket

usually

pressure only

drops

be

engine

treated

as

control

being

chargeable

evaluated

to the

accurately

The

Ap

true

components

through

7-22

shows

fluid-flow-control are

4 diameters

downstream

ated.

This

the

tapoff

nents.

The

U-tube

manometer,

pressure

and

caused

rupted

from

drop

which

straight

pipe

(a+b)=14

is

of the

diameters,

Ap that

by an unintersize

same

and

flow

length

The

condi-

any

control

components

great

virtually the

determination

able

instead

may

be already

employing

type

to obtain and

size

available.

it is

individual drop.

It is

This

can

be done

drop

the

vary

by K

flow

duct

(7-9)

its

lar

(constant

the

resistance the

rocket

to the

VALVE

fluid-How

constant range

flow

for

of Reyn-

is turbulent.

configuration,

K

the

higher

of the

component,

of size

is

were

in all

linear

only

space,

engines

are

minimum

etc.

component,

design

considerations

in table across

should

7-4. a flow-

flow-passage he

observed:

characteristic

flow

Avoid

abrupt

changes

of flow

area

(3)

Avoid

abrupt

changes

of wall

contour,

turns

in the

flow

length

of the

(5)

Provide flow

area

component

(2)

Minimize

fluid-

propellant

following

sufficient

design.

for various

drop

the

of these

geometric

of a given

of liquid

presented

be

by design

None

sizes

pressure

not the

structural

coefficients

components

depend

However,

require

various

simi-

would

influenced

necessarily

compo-

then

and

size. is

resistance

more

dimensions),

K would

of material,

available

the

to

resistance

geometrically

the

number

of the

the

of flow-control

sizes

economy

the for

cross-sectional

leading

In general,

control

(4) setup

in 2. to

K value.

of a component

sharp

test

of

open,

tends

If a series

for the

components.

same

the

coefficient

(1)-Allow

control

cross-

path

fully

a large

by component

strength,

For

7-22.-Typical

flow

size

ratio

flow-control

i-I

Figure

minimum

A smaller

Reynolds

Average

/CONTROL

com-

size.

g.

similarity /.--MANOMETER

I

the

is designed

of component

of different

upon

over

providing

considerations

/

the

of the

K is essentially

resistance

standards,

__

the

area the

independent

design

VALVE

this about

with

influenced pv 2

CONTROL

area

when

as

a higher

nents

corre-

Ap = K288---_

through

in the

area

type

coefficient

which

coefficient using

may

nearly

desir-

data

is

have

numbers, a given

the for

test

of the

passing

area

component

For

have

test

from

resistance

pressure

and

of a component

of pressure

a component

engines

conditions,

to extrapolate

calculating

of fluid-flow-

in rocket

of service

impossible

for every

number

used

variety

data

when lation

large

velocity

component

coefficient

given

olds of the

3

component.

tions. Because

flow

This

Usually

is ob-

measured

same

at the

net

psi

lb/ft

flow

sectional The

com-

method

lb/see

ponent.

the

7-22,

test

fluid,

of fluid

= characteristic

produces

component

caused

A*

gages,

data.

rate

to the

by the

fps

component,

compo-

tank

the

= flow

by the

test

by the

component,

10 diamat

weighting

of the

characteristic

pA*

disturbances

caused

repeatable

v= 144 _=

to be evalu-

of pressure

by subtracting

pressure

flow

points

and

drop

tained

the

for

chargeable

defined

in figure

=density

taps

and

component

combination

accurate

setup

Pressure

upstream

of the

minimizes

pressure

quite

test

components.

located

eters

a typical

as

shown

flow p

Figure

drop

ponent

tests.

the

= pressure

can

actual

285

where

components

turbulent.

AND VALVES

the

and

path flow

path

within

component a smooth passages

surface

finish

for the

286

DESIGN OF LIQUID PROPELLANT

TABLE Various Liquid

7-4.-Typical

Resistance

Propellant

Components

Rocket

for

main

oxidizer

of the

stage

engine.

Coefficients

Fluid-Flow-Control

ROCKET ENGINES

valve

Resistance coefficient K

Butterfly-type valves (fig, 7-33): 90 ° open ............................. 80 ° open............................. 70 ° open ............................. 60 ° open ............................. 50 ° open ............................. 40 ° open ............................. 30° open ............................. 20 ° open ............................. Ball-type valves (fig. 7-38): 90 ° open ............................. 70 ° open ............................. 50 ° open ............................. 30 ° open ............................. 20 ° open ............................ 10 ° open ............................

0.31 .41 ,77 198 5.68 15.45 44.7 124.2

Poppet-type valves (fig. 7-40): Full open Venturi-type valves (noncavitation) (fig. 7-41) ............................ Gate-type valve (fig. 7-42): Full open ............................ _Aopen .............................. ½ open ............................. _Aopen .............................. Poppet-type cheek valve (fig. 7-60) ....... Swing-gate-type check valve (fig. 7-61) .... Standard tee ........................... Standard elbow (90 °) .................... Medium sweep elbow .................... Long sweep elbow ..................... 45° elbow .............................

0.81 1.58 3.6 18.2 63 362 2.5-3.5

Liquid

oxygen

flow

Liquid

oxygen

density,

0.92 .56 .50 .50

d_/d I --tA ........................... d:/d 1:½ ........................... d2/d I :¾ ...........................

0.42 .33 .19

diameter,

Flexible

duct

actual

length,

Flexible

duct

equivalent

resistance

due

= 78 percent

will

be

further

A-I

discussed

through

in section

gpm

lb/ft in

L= 16 in

length

considering

passage

contour

characteristic

of duct

3

d=8

devi-

flow

area

area

Estimate: (a_.) The

pressure

drop

chargeable

to the

duct

The

pressure

drop

chargeable

to the

valve

Solution (a) Oxidizer The average

flexible duct flow velocity Q

is

_-

From

table

6-3,

0.277

× 10 -_

the

Dvp --_

Use

8x _

/z

duct.

chargeable

eq.

7-6).

in the

duet

0.00005 8 12

From

figure

equation to the

=2"94x

107

projection

size

• of

roughness

Substitute

(L e + L) into

flow

x 10 -s

roughness

or a relative

D

(see

of the

oxygen × 10-7

79"4 x 71"38 0.1282

a surface

0.00005

fps

= 0.277

lb/ft-sec

number

duct

of liquid

thus/1

10 -3

Reynolds

flexible

- 79.4

viscosity

lb-sec/in2;

x4636.8-0.1282x The

in the

12 420 d2 =3.12xTrx16

3.12x

f: 0.0112. flowing

to flow

ation, Le = 6 d Main oxidizer valve

Re

of fluids

Q = 12 420

inside

for the

orifices

of the

p = 71.38

duct

v= 0.18 120 5.6 24 2to 4 1 to 2.5 18 .90 .75 .60 .42

rate,

Flexible

0.8-I.5

Sudden enlargement: d_/d 2= ¼ ........................... d_/d2='_ ........................... dl/d2=_A ........................... Ordinary entrance ...................... Sudden contraction:

characteristics

type)

Engines

Component description

The

(butterfly

the

0 000075

7-20,

(7-7).

oxidizer

friction

equivalent The flexible

factor, total

pressure duct

length drop then

is

7.10. f(Le Sample The oxidizer

Calculation following pump

Ap =-

(7-2) design

discharge

data

are

flexible

given duct

for the and

the

+ b)pv

2

288 gd

_ 0.0112

(6 x 8 + 16) x 71.38

288 x 32.2 x 8

x (79.4)

2 = 4.34

psi

DESIGN OF CONTROLSAND VALVES

(b_._)gain oxidizer The characteristic

valve velocity

79.4 V:o.--._18= 101.6 From table

of the valve

fps

7-4, the resistance

butterfly valves K=0.31. Substitute equation (7-9) to obtain the pressure able to the main oxidizer valve:

Pv2 0"31×71'38×(101'6)2 Ap = g-_8_g 288 × 32.2

Control Fluid Pressure

coefficient

for

this into drop charge-

= 24.65

psi

Level

The working pressure level and the temperature of compressible fluid-flow-control system are important factors, since both govern the density of the fluid. Means of compensation for changes of pressure in a compressible fluid control system must always be provided. With an incompressible fluid, the pressure has relatively little influence on density. The working pressure level of the fluid determines the selection of the structural design of the components as well as of the sealing methods, especially for dynamic seals. Special provisions are often made to meet the stringent requirements in high-pressure applications. For example, the cutoff events in a high-pressure turbopump-feed engine system may be sequenced so that turbine power is cut first; thus the main propellant valves are not required to shut off against the high main-stage discharge pressures.

Fluid-Flow

Velocity

The requirements for smooth component-flowpassage contours are more critical with controls for compressible, or low-density, fluids such as hydrogen than for incompressible fluids, because their design flow velocities usually are much higher than those of the denser liquids. Also, in general the design trend for high-thrust, highpressure engine systems is toward smaller propellant duct and valve sizes, and consequently toward higher flow velocities (over 100 fps). An important consideration in the design of high velocity flow-control components is the high-impact loading imposed upon the control

287

surfaces by the fluid stream. This is especially acute with the higher density liquids. To obtain reliable control performance characteristics with liquids at high velocities, the control components subject to impact loading must be designed to withstand the stresses involved. Also, they should be contoured so as to maintain small impingement angles with the fluid stream and to keep inpact forces to a minimum.

Fluid-Flow

Temperature

Temperature is an import_tnt consideration for the design of fluid-flow controls. This is especially true if the controls are for fluids at temperatures in excess of, or far below, norn_al ambient. In liquid propellant rocket engines, fluid-flow controls may have to handle hot gases at temperatures up to about 1700 ° F. Example: the control of a turbine working fluid. Hot liquids need not be considered, since none of the liquid propellants have sufficiently low vapor pressures to permit handling at high temperatures. Ability to operate at elevated temperatures without any form of lubrication is a prime objective in the mechanical design of fluid-flow control. This can be accomplished by using bearings of either extremely hard, wear-resistant alloys, such as stellite and sintered carbides (high loading condition), or relatively soft materials such as graphite (low loading condition). Bearings are usually subject to compression loads only and are therefore not subject to failure if the materials used are of low ductility. For structural members not subject to wear or bearing loads, conventional high-temperature alloys such as stainless steels and other nickel-base alloys may be used. For static and dynamic seals, metal gaskets and bellows, carbon or graphite face seals, and labyrinth-type seals are suitable at high temperatures. At the other end of the scale, liquid propellant rocket engine controls may see extremely low-temperature levels, such as in liquid hydrogen service (-425 ° F). Hero, two principal conditions must be considered: (1) The physical characteristics of the fluids which at these low temperatures may affect control performance; and (2) the physical characteristics of the materials from which the control components are made and

•

288

[

.

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

which may affect the operation and, thus, performance characteristics of the control devices.

the

Many of the cryogenic fluids, i.e., liquefied gases, experience somewhat unpredictable phase changes (two-phase conditions) for relatively small temperature changes. No serious difficulties need to be expected, however, if the heattransfer rate from components in critical control areas is low enough to prevent vaporization of the liquid. This is particularly important in liquid hydrogen service, where insulation may pose difficult design problems. At any rate, except for viscosity changes, nearly all liquids exhibit more stable physical characteristics with large temperature variations, within the range between their freezing and boiling points, than do gases if the temperature range reaches to their liquefaction temperatures. The construction materials for fluid-control components for low-temperature applications must be especially carefully selected. Practically every metal undergoes irregular phase changes at low temperatures which may seriously affect its physical properties. While the strength of metals generally increases with a decrease in temperature, further temperature decrease beyond certain limits may result in a decrease in strength. Many metals also become brittle at very low temperatures. Most of the aluminum alloys and the 300-series stainless steels exhibit much better stability at temperatures in the cryogenic range than do others. Elastomers such as Teflon, Kel-F, and Mylar, when used for sealing purposes, exhibit satisfactory mechanical characteristics at extremely low temperatures. Teflon-coated surfaces additionally have good anti-icing characteristics. For further detail on materials, see chapter II. Fluid-flow-control components for operation at cryogenic temperatures should be designed to be free of external icing effects. In addition to insulation, moisture-preventing purges should be provided internally in critical areas such as bearing interfaces. Also, actuators and/or bearings may require heating.

Rate

•

of Response

in Fluid-Flow

Controls

Response rate is an important design eration in any control system. Basically,

considthe

limiting factors governing response rate are (1) the speed with which signals can be transmitted, and (2) the mass/force ratio or its function, the inertia/force ratio of the main control organ. In many fluid-control systems the controlled fluid is used to transmit the sensed signal, In others, part of the sensing link employs electrical or mechanical means. However, in most cases, part or all of the sensing loop utilizes an impulse generated by a pressure change. This impulse is transmitted at the speed of sound in the fluid. As a typical example, the velocity of sound in water is five times that in air; accordingly, a control signal would be transmitted five times faster in water. The actuators for most fluid-flow-control mechanisms use pistons or diaphragms, powered by fluid pressure which, in turn, is regulated by some form of pilot valve. If suitable, the controlled fluid may be used as the actuating fluid. The response and flow capacity of the pilot valve, the effective area of actuator piston or diaphragm, and the actuating fluid pressure level influence directly the response rate of the control mechanism for given mass inertia and frictional or other resistances. To satisfy attain stable

certain operating conditions control it sometimes becomes

and to nec-

essary to introduce simple damping devices. In most control systems, stability is inversely proportional to sensitivity or response rate. Thus, the design of a fluid-flow-control system should reflect a realistic balance between sensitivity or response rate, control accuracy, and system stability. Figure 7-23 illustrates the schematic of a typical piston-type actuator for fluid-flow-control devices. The piston when actuated moves against the spring in the direction of the arrow. The basic correlation between the response rate or acceleration of the piston, and other operational parameters, can be expressed by Map -Alp g

1-A2p2-

Fr- Ft-

Fs-

Cx (7-10)

where M :effective piston, that of nected all the

mass accelerated by the actuator lb. It consists of piston mass, moving parts mechanically conto the piston, and of the mass of fluid columns in the system

DESIGN OF CONTROLS AND VALVES

ACTUATING SUPPLY

FLUID LINE

i SPRING

tons,

shafts,

Here,

too,

289

and

rods;

tant

design

fied

into

seals

is one

and

seats.

most

Seals

can

impor-

be

for medium-temperature

(-60 ° to 400 a F), ° F),

for valve

of the

considerations. those

to -425

and

temperature

classi-

service

low-temperature

service

(-60

high-temperature

service

(400 _ F

°

0 FI BODY

and

up).

the

materials

extent

The

on

service

involved. VSNT _-__!I!!

///,

seals

outstanding

p,sTo; N"I

Figure

7-23.-Schematic

they

of the

piston,

ft/sec

of the

piston

actuating

A s =area

of the

piston

vent

p_ =actuating

pressure,

side,

psia.

This

less

supply-line

pressure

on

rate

the

depends

the

acceleration P2 =vent

psia.

pressure,

plus also

function Fr = resistance

is

fluid

the

tile

drop

also

may

tion ap Ff = friction forces

is

the

vent-line

depends

sliding

drop

(again

function,

be a function

spring

=spring

rate,

force,

x

=distance

traveled

initial

position,

lb (at

a

practices seals

used

surfaces,

Design

x = 0) piston

from

its

DYNAMIC

SEALS

FOR

may

the as

Apart treated seals nents: ing)

from

in chapter are

required

seals cylindrical

the

static

seals,

which

IX, two

basic

types

will

elements

as

provide

backup

compo-

mended

(reciprocating such

will

of dynamic

for fluid-flow-control

for moving

be

actuator

and

rotatpis-

sures

will as

pressures

S00 above

is

as

psi

and

psi.

A

(table

7-25

In is

conusually

In addition,

seals

hard7-5).

diameter

designs

for static

and

O-ring

installation

in figure

O-ring

1500

safely. and

Past guide.

and

pressure maximum

section

space

permit.

seal

of compound

the

extrusion

nominal

shown

O-ring

of fluid

be tolerated

for dynamic over

pressure

of clearance

a useful

rings

leaving

the

can

O-ring

large

siderations

7-25),

determine

O-ring

O-ring

(fig.

will

prevent

is

O-rings

choice

combination

E that

failure

influences the

pres-

diametral

to the

after

it affects

combination

fluid

of seal

adjacent

seal

considsummarized

and

pressure)

pressure

The

chosen

COMPONENTS

gap

hardness

general,

cause

appli-

O-ring

are

between

deformation

because

ness

FLUID-FLOW-CONTROL

seals

hardness

under

Fluid

proper

OF

clearance (when

clearance

for

O-ring

correlations

static

design

frequent

plex and solution.

computer

and

of diametral-squeeze-type

chosen

success.

a typical

Important

compound

extrusion

design

in

as

design

recommended

dynamic shows

clearance.-A

reduced.

proper

represent

the

the

applied parts

for diametral-squeeze-type

seat.

hardness.

7.7 DESIGN

imper-

widely

cylindrical to assure

7-5

for a valve

O-ring

been

be observed

for typical

a permanent

a high-speed

minor part.

However,

table

7-26

1. sure,

have

seats.

Figure

groove

lb/in by the

The

is that

for Medium-Temperature

for moving

erations for dynamic as follows:

lb,

of accelera-

and

O-ring

Since the relations between p_, P2, Fr, and ap are not linear, equation (7-10) may become comrequire

must 7-24

design

into

C

seals

despite

O-rings

for valve

cations.

ambient

rate

ap) control

as

Figure

of

lb

Fs = initial

or elas-

possible.

mating

Seals

seals

techniques

which

pressure

on flow

(seals,

type

wherever

or the

of Dynamic

dynamic

well

source,

(a function

This

of acceleration force of the

which

etc.),

flow

as

ap)

pressure,

which

actuating

of fluid

of these

seal

Elastomeric

in s

in 2

at the

and

2

side,

pressure

to a large

nonmetallic

satisfactorily

in the

Design Services

A_ --area

used

and

is based

piston-type

actuator. ap = acceleration

soft

are

function

configurations

conditions

advantage

fections

of a typical

of the seals

Generally,

tomeric

A.

selection for these

Teflon are

recom-

at sealing O-ring

seals

presat

illtBl #: 29O

DESIGN OF LIQUID PROPELLANT

ROCKET ENGINES

STATIC

SEALING\ _ \BREAKCORNERS,

¢,-_ I'-: _b

"1 -,J

?o APPROX. o.oos' D

"="

. _L__L._4___ a_-_--L

D-GROOVE U:NGTH

l .-f

"-

CENTERLINE OF THE PISTON ACTUATOR

Figure 2. the

7-24.-Diametral-squeeze-type

Surface

sliding

O-ring

finish

requirements.-The

surfaces

seals

O-ring

in contact

should

be

as

finish

with

smooth

seals of

dynamic

as

that

longer

life

a finish than

ishes.

Codirectional

honing,

have

finish

better

surface and

again

slippery

proven

type

possible.

after

be hard

finished. surface

scratching.

The that

yields

as

best

sliding

an initial

plating

resists

the

plated

provides

of 60

7-5.-Recommended

rms

or bet-

Design

Practice

[See fig. 7-15 for explanation O-ring nominal section diameter

O-ring section diameter

1/16 3/32 1/8 3/16 1/4

0.070± 0.003 0.103 *_0.003 0.139 *_0.004 0.210 *_0.005 0.275 *_0.006

Diametral

squeeze, rain

Dynamic 0.010 .010 .012 .017 .029 Fluid

= +0.000, : +0.000,

-0.001. -0.005.

and

with N,

selection

the frictions

trade

increase

can

only

Accurate

design.

elastomer

compounds

names

such

as

Teflon,

of an O-ring

all dimensions

with

be obtained

compounds,-A

Viton,

is

diametral

of O-ring

great is avail-

Silicone and

compound

for Diametral-Squeeze-Type

of dimensions;

seal friction.

of temperature.

frictions for a given

Butyl,

seals.-The

running

hardness,

with

seals.

O-ring

O-ring

of O-ring

Buna

times

decrease

Selection

able,

O-ring

running

of O-ring

4.

Kel-F. and its

O-Ring

rubber, The physical

Seals

in inches] 2 x E-

C-glandwidth

aDynamic

0.015 .017 .022 .032 .049

0.057 .090 .123 .188 .240

pressure

and

experimentally

Static

(3-1000 psi 1000-2000 psi 2000 psi and higher aTolerance bTolerance

and

in contact

of a dynamic

three

pressure,

variety

a hard, wear,

applications.

O-ring

of dynamic

about

squeeze,

For

static

for surfaces

friction

usually

values

seal.

corrosion,

finish

by

finishing,

or nickel

Friction

breakaway

fluid

surface

and

diametral-squeeze-type

Breakaway

fin-

produced

to be the

dynamic

is recommended

3.

a

or smoother

chrome

A microinch

TABLE

range

of dynamic

results,

could

this

rougher patterns,

been

for any

still

within

either

ter static

They should be ground, honed, or polished to a microinch finish of 8 to 10 rms. It has been found

in typical

D-

bstatic 0.052 .083 .113 .173 .220

O-ring compound hardness 70 Shore "A" Durometer 80 Shore "A" Durometer 90 Shore "A" Durometer

groove length

R-

radius, rain

diametral clearance max

3/32 9/64 3/16 9,f32 3/8

1/64 1/64 1/32 3/64 1/16

0.005 .005 .006 .007 .008

291

DESIGN OF CONTROLS AND VALVES

_AOIAL

-,-x

_"

C k F.A,_I* N C _ __]_

_PRESSU_E

__

P_ESSURE

_"

7-27,

7-28,

dynamic valve

EXTRUSION UNDER

t"igure

OF

(_RING

RJNGS

PRESSURE

o[

O-ring

of the

under

backup

pressure

seats.

the

application

rings.

(dynamic 5.

of seal

has

and

been

applied

at temperatures

are

also

pressures

as low

reasonably

as

effective

molecular-weight

gases

such

over -425

when as

2000

° F.

psi,

They

sealing

helium

low-

and

'Z:,;_ ...........

°"' _°

I 1710 t

010

installation

of O-rings

....,oo,I

dur-

1, ...../

\ 1

important

k t, "-/->' 1 ;

L

chamfers

should be provided on all edges and in contact with O-rings to minimize the

possibility

of cutting

or scratching

during

the /

assembly

process.

L

6. O-ring seals for valve seazs.-O-rings be applied effectively as seals for valve

can seats

(fig.

absorbs

7-26). loads

The

resiliency

and

seals

of the

tightly

O-ring

at all pressures,

even when some dirt and grit are present system.

type

as parts

o[ diametral-squeeze-type

ing component assembly is extremely to assure an effective seal. Generous

shock

effectively

cylindrical

or static).

seals.-Proper

or radii corners

used

hydrogen.

Installation

O-ring

are

at sealing

and

properties (furnished by the producer) is based on operating conditions such as type of fluid or propellant, pressure, temperature, and type of seal

This

successfully

diamelraI-squeezeand

7-29)

for moving

PREVENT

EXTRUSION

7-25.-Extrusion

type

TO

and

seals

One

design

problem

is

k _IPSEALS 121 _(_QLIIRIEO

_TE

--A_A_ "_ _ _ °PIER

UNLESS aT_E_SE SPECJFP[I]I F_NISH SUtteES ANC} t_ SEALS

_

to

Figure

7-27.-Typical valve

rotating

lip-seal

actuator

shalt.

_

T_E

CONTACT

O_l i'LL ._AkS

WiTH

TO BE 12 RM$

OETTER.

in the

to prevent

_.

ALL

THE OR

NICENSIOhS

_H _NC_ES

design

[or

the

O-ring from being blown out of the groove. This can be prevented by providing a dovetail O-ring groove in a two-piece valve poppet (see fig. 7-26).

Design Services

of Dynamic

Seals

for Low-Temperature

For cryogenic or low-temperature services, lip-type seals made of elastomer sheets (figs.

VALVE

STEM

Figure VALVE

PRESSURE

'

POPPET

7-28.-Typical

[or butterfly-type

VALVE

_

seat

lip-seal

design

valves.

OUStNG

The basic employ ,_AL

valve

PRESSURE

I

pressure

RETAINING

the

design fluid

principle

pressure

at the sealing

of lip seals

to increase

surfaces.

is to

the

contact

Due to their

RING

lip Figure

7-26.-Typical design

valve

for poppet-type

seat valves.

O-ring

seal

configuration,

is maintained Design

the

resilience

of these

seals

even at very low temperatures.

considerations

for lip

seals

are

similar

292

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

-Acrvz.

to those for O-ring seals. The design approaches can best be illustrated by examples. Figure 7-27 shows a typical valve shaft rotating lip seal arrangement, including dimensions and surface finishes, for liquid oxygen and hydrogen service. Figure 7-28 presents the valve seat lip seal of a butterfly valve for use with the same liquids. Valve seat O-ring seals (fig. 7-26) made of Kel-F have also been successfully applied in poppettype valves for liquid oxygen. The design of lip seals for piston-type actuators using lowtemperature helium gas as the actuating fluid is shown in figure 7-29. In liquid hydrogen service, metallic bellows (as shown in fig. 7-30) have been used to a great extent to achieve positive dynamic sealing. However, pressure levels and available space impose limitations on their application.

/' _o_

( CLOS_N& )

The metallic

Seals

bellows

FOR

SEALS

_.,:

.

',_

../'

;b"_

_L_tE_

_

.

}

001NV_ _.&LIN*

5t._f_3t5

LBO_H

_LVE

AN0

,_v_t

) L_PPED

N

8-10

B£LLOWS USEO AS Ty_P'_ D'yINiAIMI_;

(fig. 7-30) is most fre-

DOUBLE

LIP

i

ASEhqI._

Figure 7-30.-Metallic bellows used as reciprocating-type dynamic seals in a typical poppet valve for high- and low-temperature services.

for High-Temperature

quently used as reciprocating-type dynamic seals for high-temperature services. Two types of metallic bellows are distinguished: the hydraulicformed and the multidisk welded type. The former is made of one to three plys of sheet metal and is designed for all pressure ranges. The latter is for relatively low-pressure services and for high flexibility. A metallic bellows of any type behaves, in part, like a helical spring. The spring rate (lb/in of.movement) is a direct function of the

DOUBLE

.-T_P,Ca_ WEL._EO .N_N_ S,{'rUJXOm I,LA'II KLLDW$

W:

WELI)ED TYPE RIEC_PRI_C,ATING

Design of Dynamic Services

_

}_

_%ulO

,_

LIP

mum.

Generally, bellows design data, such as stock size, allowable working pressure, spring rate, materials and service temperature, are supplied by the manufacturers. Important design considerations are discussed in the following:

SEALS

ACTUATOR

elastic modulus, and of approximately the square of the thickness of the material. It is also a function of the outside-inside diameters and of the number of convolutions and their curvature. For maximum flexibility (inches of stroke/lb of load), a minimum inside diameter combined with a maximum outside diameter should be used. Also, material thickness (within stress limitations) and modulus of elasticity should be mini-

PISTON

/

1. Applicalion of pressure.-When a bellows is subjected to a differential pressure between interior and exterior, it is preferable to apply the

VALVE

[_ __-t__

_

_j;! _

ACTUATOR

i'_.>_AC

T UATOR

ROD

PISTON

Figure 7-29.-Lip seals for piston-type actuators. Double lip seals seal pressures both ways.

higher pressure to the exterior. This reduces stress, and permits higher pressures and longer life for a given design. 2. Provision of mechanical stops.-These should always be provided to prevent extension of the bellows beyond its permissible extended length and compression beyond its "bottomed" height.

293

DESIGN OF CONTROLS AND VALVES

3.

Selection

lows

or materials.-Selecting

material

patibility

should

or corrosion

temperature Some

and

high-temperature

steels,

Monel,

proven

suitable.

4. when yields

spring

such

and

Hastelloy

Inconel,

EHective

area.-This by a change

actual

is

can he approximated

that

as

stainless

volume.

nate

design

ring

to the

effective

× (inside

area

by

area

diameter

= 0.1963 +outside

diameter)

5. End attachment.-Typical the end attachment figure 7-30.

welded

2

(7-11)

joints for

of bellows are illustrated in

for low-temperature

was

up

shaft is

sealing

to 700

(fig.

as

psia.

has

7-31).

seats

basic

alterseal

which

is

the

shaft

Here, by the

flexibility

for high-temperature by metal-to-metal

in figures two

An graphite

bellows

achieved

shown

success-

1200 ° to

a flat-face

of valve

contact,

7-30

and

7-32.

requirements.

a finish

of 10 nns

or better

sealing

surfaces.

Secondly,

loading

must

be applied

deformation

from

compensated

is usually

design

operated

ranging

of a metallic

services

This

Firstly,

is required

for the

a high-enough

to create

unit

a compensating

of the sealing surfaces

and

to

achieve the intimate contact required to overcome

Silver brazing and soft soldering

can also be employed ices.

end

to the

The

length,

seal Bellow

which

is to attach

misalinement of the bellows.

which,

This

valve

at temperatures

welded

B have area

throttle

fully

1S00 ° F at pressures

operating

in bellows

displaced

com-

characteristics.

alloys

multiplied the

by fluid

considerations,

ranges,

gas

of bel-

be governed

manufacturing

tolerances, distortion of the

serv-

A typical design of a rotating-type dynamic seal for high-temperature services is illustrated in figure 7-31.

The

dynamic

PRESSURE HOT GAS

HOUSING VALVE

_

sealing is achieved

through the spherical mating surfaces between the graphite seal ring and the steel shaft collar. The

contact force of the sealing surfaces is

maintained

by the shaft thrust spring. Any

alinement between

shaft is compensated and side movement arrangement

mis-

the thrust bearing and the by the spherical seal face of the seal ring. This seal

has been applied to a turbine hot

SEAL

"X

O

SURFACES

ROTATE

FINISH tO OR BETTER

{a)

RMS

TO OPEN VALVE

CONVOLUTED HOT

_ •. _

/-

_..

_

\ BUTTERFLY

R_NG

GAS

VALVE

SHUTOFF

TYPE

VALVE

SEAT

BUTTERFLY

SEAL

SPRING

GAS

DISK

VALVE

_,

_

VALVE

sto,,

/

_b} S(AL _s, aN

typical

hot

rotating-type

gas

throttle

dynamic

POPPET

/

|

/ ......

7-3I.-Turbine

_N

_1_?,_,_ _ _,,_ HOT

Figure

USED

VALVE

PRESSURE

VALV¢_

GATE

VALVESHAFT

valve seals.

with

POSITIVE

SEAL

STOP

IO RMS

SPRING HOT

Figure

DISK GAS

TYPE

SHUTOFF

7-32.-Valve ture

VALVE

SEAT

POPPET

seat seals services.

HOUSING

SURFACES

SEAL

OR

FINISH

BETTER

USED

IN

A

VALVE

for

high-tempera-

A

_1

mi

_rx

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

294

valve parts due to temperature, internal stress, and mechanical loading of the mating parts. Figure 7-32(a) shows a convoluted-ring-type valve seat seal used in a hot gas shutoff butterfly valve. Depending upon the specific application, the convoluted ring may be made of hightemperature alloys such as Inconel-718. The rings effect a leakproof seal in the closed position, since the upstream fluid pressure tends to expand the convolute and produces a high contact unit force at the sealing surfaces. The curvature of the convolute ring tends to maintain a continuous contact with the valve seat. Figure 7-32(b) presents valve seat seal

a metallic-spring-disk-type used in a hot gas shutoff

poppet

valve. Again, the upstream gas pressure produces a high contact unit load on the sealing surfaces. The valve seat has a curved contour which effects a continuous contact with the flat face of the seal disk.

Sealing

Specifications

The degree of sealing (or the allowable leak rate) is a very important specification which will dictate the type of seal to be selected for a specific fluid-flow-control component design. The basic reference for leak rates is Specification MIL-S-8484. It states that a Grade A seal, the highest quality seal, shall have a leakage rate not to exceed 1 standard cubic centimeter

of

air/year/inch of seal at a pressure differential of 1 atmosphere. This corresponds to a leakage rate of 3.171 × 10 -s cc/sec/inch of seal. It is a design assumption that any seal leak rate below or equal to this value is considered zero leakage. For many applications, higher leak rates are permissible. For instance, a check valve may be specified with a leak rate of 5 scim's (standard cubic inches of gas per minute). This is still a relatively tight specification.

7.8 DESIGN

OF

PROPELLANT

VALVES

Propellant valves are used to initiate and terminate propellant flows to main thrust chambers and gas generators. They are usually openclosed, two-position, normally-closed valves. To meet specific sequencing requirements, other designs may provide for an intermediate opening position. For thrust-throttle or mixture-ratio-

control purposes, ability for continu,_usly variable opening position may be required. In addition to propellant compatibility and structural integrity, prime design considerations for propellant valves are: (1) No leakage of propellant through the valve when closed (2) Proper actuating time during opening and closing in accordance with the requirements of the control system (3) Minimum pressure drop A great variety of propellant valve types is available. Each design has certain characteristics which make it suitable for a specific application. Frequently used propellant valves, classified according to their design configurations, are: (1) Butterfly valves (2) Ball valves (3) Poppet valves (4) Venturi valves (5) Gate valves (6) Needle valves Butterfly-Type

Propellant

Valves

The butterfly valve is one of the most widely used propellant valve types in large liquid propellant rocket engines. It has established a reliable operational record in LO2/RP-1, LO2/ LH 2, storable, and other liquid propellant services. Existing butterfly valve designs range from 2 to 17 inches nominal diameter, for use at propellant pressures from 20 to over 1500 psia. With improvements in sealing and structural details, successful designs for higher capacities and propellant pressures are certain to be achieved. Fignre 7-33 presents a typical butterfly valve design. Sealing is provided by a lip seal, which engages a spherical surface on the valve gate, similar to figure 7-28. The valve gate pivots on the valve shaft, the axis of which passes through the geometric center of the spherical sealing surface. In most designs, the valve gate rotates 90 ° from the closed to the fully opened position. The valve is operated by a piston-type actuator, through a connecting link and shaft crank arm. Lip seals are used as dynamic seals for the rotating valve shaft (fig. 7-27). The actuating power is furnished either by noncryogenic propellant pressure, or by an inert gas supply, and

295

DESIGN OF CONTROLS AND VALVES

plsroN

_

O"Rr _'_pE

_-

/

/

-_

--

\

++

/

SEe:tON

:

A

i'

_-AT FL;LLYOP_NEO PO_ITIO"+, _ i

+ ...... ; .... • I,o+,,,o.......... + ......,.+o ....... -

\

A--A\

u_

L_.......o X__

+ I

.......

+,

+-

r---,_

valve

+ "-:"+-._

:

butterfly-type in the

piston. are

a pilot on the

Except

7-34

shows

gen

valve

booster

steels,

a 4-inch, used

of the

pins

most

of the

the

which

Butterfly

valves

have

relatively

flow.

They

are

to service. area

which

They can

have

be

valve

booster

engines.

liquid

A butterfly

smooth

fluid-flow

valve-gate

s

as

a throttle

(see

fig.

7-33

low

compact, a high

parts

oxy-

local

flow

A* = characteristic ds =inside in

area

diameter

characteris-

as

the

A-1

freezing.

The

butterfly

valve

(%12)

the

valve

position, Values area

nominal about

references) valve,

valve

seat

means

in _ lip

for A* (duct diameter,

87 percent

gate

area

at the

fully

open

in _ range

area

from

= v/4 in)

65 percent

dn 2, where

on a 2-inch

of the

duct

area

size

of the valve,

on a 12-inch

pressure.

linkage

between fuel

engine. attached ates

d_ = valve

the

During to the

shaft

position

of the

valve

the

engine

control

may

may

be

from of the

provides system open

indicators

for

system

RP-1

be normally

or nor-

may be

added;

be

accomplished

by

squib,

rather

by

main

7-35

valve

sequence

of the stroke, valve valve is also

indication gate.

than

illustrates

oxidizer

a potentiometer for continuous

fluid

a heater

7-33

the opening main oxidizer fuel

in table

arrangement

Figure the

sequence

_he igniter

Frequently, to

valve

opening

listed

in figure

of a pyrotechnic

pneumatic seal,

may

fluid-

actuating

to keep

position

of the

the

actuator-valve

valve

closed;

are

engine,

actuator shown

such

Typical

in a LO2/RP-1

for specific

needs: mally

as

stage

toward

effects

K at various

valve,

of used

tendency

valve

is used

at the

range when

cavitation.

oxygen

a relatively

Thus, little

coefficients

RP-1

required

igniter

Ag=projected

duct

of the

ICBM

a wide

adverse

of a butterfly

liquid

such and

it has

attendant

resistance

When

resistlight,

_-Ag

of the

liquid

Atlas

maintains over

positions.

propellant

positions 7-4.

ICBM

as

for dimension

main

Rocketdyne

valve

valve, with

closing where

on

stream

angular

flexibility A*-_-d -4

butterfly-type

used

valve.

as

Figure Atlas

expressed

7-34.-Four-inch

oxygen

the

to fluid

easy

Figure

turbulence

are

other

forgings.

Rocketdyne

is

actuator

which

butterfly-type,

on

shown

engine.

ance tic

and

of aluminum-alloy

,......

by a spring

side

for shaft

_

position).

valve

closed

closing

of stainless made

The

i"

propellant

closed

valve.

to be normally

installed

made

r --:-_ .... ---_

(shown

by

+,.<

_ +

design

designed is

,

7-33.-Typical

controlled

%A z,

I+.'-C

//.

Figure

•

valve A-1

the and

the

stage

the cam shaft actuto open. attached of the

to angular

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

296

_._tN

OXJDtZEk

OPE*WtNG P=E_J_E

VALV_

SIDE ADTIJATING. APPUED M,&pN OXIDIZ_I VALVE iN FULLY C_=ENED POSITION

FWID

AC IIJAIIN G

_CLOSlNG)

FtUIP

FUEL PIIS_U_! E_G_NE

INLET PORT AC_UMING (C_ENI*WG)

r_ow

CO_WIItOL

°

_, VALVI

POIATE TO J O_EN VALVt VALVl

SH_.= T --_

IGNiIEI FUEL SEOU_NC_ VALV_

IGNI_EI /,ND

C_mrDC_

IGNmON

MONIIO_

Figure 7-35.-Mechanically between the main oxidizer fuel

sequence

valve

VAtVt

linked arrangement valve and the igniter

of the A-I

stage

engine.

BUTTERFLYVALVE GATE

./

AXISOF VALVE / SHAFT

FLOW (a) DIRECTION

The amount of torque required to turn valve shaft and gate is determined by the summation of hydraulic and friction torques. Hydraulic torque is the unbalance of forces on the valve gate caused by the flow of fluid around it. If the axis of the valve shaft is located as shown in figure 7-36(a), the fluid striking the gate portion protruding farthest upstream is deflected more than that at a point near the other end of the gate. This produces an unbalanced force which tends to close the gate. Offsetting the valve gate as shown in figure 7-36(b) would further increase the closing torque, because the fluid velocity rises as it approaches the downstream side opening. Consequently, the resulting low-pressure, area tends to increase the unbalance in the closing direction. For this reason, butterfly valves are usually designed offset as shown in figure 7-36(c) (also see fig. 7-33). This produces a fluid velocity effect tending to ease opening of the gate, because of the lower net closing hydraulic torque acting on the valve gate. Nevertheless, the net hydraulic torque will still be acting in the closing direction for most angular gate positions (9°-80°), unless the valve gate is further offset. Friction torque always opposes rotation. For most operational valve designs

FARTHESTUPSTREAM PROTRUDING

To = Tt+ Th

(7-13)

T c = Tf-

(7-14)

Th

where To Tc Tf Th

FLOW (b) DIRECTION BUTTERFLY VALVEGATE

AXISOF VALVESHAFT ./

= required opening torque, in-lb = required closing torque, in-lb = friction torque, in-lb = hydraulic torque, in-lb (assumed the closing direction)

to act in

The friction torque Tt varies with the pressure differential across the valve gate, and with the valve gate projected area which is a function of gate angular position. Friction torque can be estimated by

(c) FLOW DIRECTION _

T! = Ktrsfmds2hp

____

(7-15)

\

BUTTERFLYVALVEGATE Figure ?-36.-Various axis with respect

locations to butterfly

of valve shaft valve gate.

where K[ = friction torque coefficient, which is a function of gate angular position (to be determined experimentally)

297

DESIGN OF CONTROLS AND VALVES

rs

=radius tion,

of valve in

fm = coefficient

shaft;

steel

shaft)

between

for aluminum 0.05

sec-

estimated

shaft journal

for needle

and

ator

has

and

all

seals.

bearing

and

of valve

differential

seat

across

lip the

seal,

torque

Th may

gate,

Kh ---hydraulic

torque

(7-16)

a function (to Figure and

7-37

closing

tions will

torques

type)

gate

of the

actuator

to three

the

forces

Opening

actu-

valves and

of

are

closing

milliseconds.

stage

diameter

needle Data

Test

experimental

oxidizer

valve

data

(butterfly

engine.

shaft

at bearing

of the

valve

section,

in seat

lip

seal,

in

Coefficient

valve

and

main

of valve

ds -- 7.7

posi-

(7-3)

of friction

between

bearing,

f m =0.05

angular deg

Ap,

shaft

and

maximum

REQUIRED OPENING (To=T f +T h )

Valve gate position,

TORQUE

5 ..............

/FRICTION

friction

Data

Inside

In actual

of a butterfly

times

static

design

A-1

rs=0.8

opening

angular

valve.

the

20 to 200

for the

Radius

position

of required

butterfly the

two

is

experimentally)

plots versus

which

angular

be determined

practice, provide

of gate

shows

for a typical

design

coefficient,

In addi-

stroke,

propellant

acting.

from

following

given

Design where

fast

torques.

opening

the

Calculation

The

by

3Ap

closing

Butterfly-type

range

Sample

are Th = Khds

of the

in

valve

be estimated

and

start

to overcome

relatively

psi Hydraulic

opening

at the

times

diameter

Ap=pressure

bearing

tion,

(0.20

steel

= inside

at the

of friction

bearing

ds

shaft

TORQUE

(Tf)

W D 0 re" 0 I--

15

..............

40

..............

85

..............

Determine torques T,

ORQUE

0°

(Th)

90 °

tions

the at the

of the

psi

K[

1058

0.78

1.11 × 10 -a

769

0.78

2.55x

87.5

1.57

12.50×10

25

3.61

required

opening

10 -_ -_

-11.64×10

and

5° , 15 ° , 40 ° , and

valve

Kh

-3

closing

85 ° angular

posi-

gate.

Solution

OPENINg" From

equation TI=

REQUIRED '

=

--

CLOSING

(7-15),

the

friction

torques

Klrsfmds2.Ap

at 5°:

T[=O.78xO.8×O.O5×(7.7)2x1058 = 1960

in-lb

at

T[=0.78×0.8×0.05×(7.7)

2×769 = 1425

in-lb

2×87.5

TORQUE 15°:

)

o

at 40°:

T[=l.57x0.8×0.05×(7.7)

o

at85°:

T[=3.61×0.Sx0.05×(7.7)

-'- 326

in-lb

I--

= 214 GATE

0=

Figure ing

2×25

ANGULAR POSITION _ CLOSING

7-37.-Typical torques

butterfly

versus valve.

required gate

/

90 °

opening

angular

From

and

position

equation

(7-16),

the

hydraulic

in-lb

torques

Th = KhdsSAp at 5°:

Th = 1.11 x 10-3

at 15°:

Th;2.55x10

x (7.7)3

closfora

x 1058 =

-3x(7.7)

535

in-lb

895

in-lb

3×769 =

DESIGN OF LIQUID PROPELLANT

298

at 40°:

Th = 12.50x

10 -3 × (7.7) 3 x 87.5 = 500

at 85°:

Th=-11.64×10-3×(7.7)3x25

in-lb

equation

(7-13),

the

aCT_TOa

J_L =-133

From

ROCKET ENGINES

required

in-lb

opening

torques To = TI+

Th

at 5°:

To=1960+535

=2495in-lb

at15°:

To=1425+895

=2320in-lb

at 40¢:

To=326+500

=

at 85°:

To=214+(-133)=

From

equation

826in-lb 81in-lb

(7-14),

the

required

closing

torques Tc = (Ttat 5°:

Th)

Tc=1960-535

Figure

=1425in-lb

at l5°:

Tc=1425-895

=

at40°:

Tc=326-500

=-174in-lb

at 85°:

Tc=214-(-133)

= 347in-lb

7-38.-Typical

530in-lb

assembly

consists

metallic Ball-Type

Propellant

The

major

Valves

advantage drop,

since

stricted

flow.

Its

fluid

tural

soundness record

propellant

applications,

generators

as

bers

(up

inal

diameter

well

space

envelope

eters,

ball weight

ever,

for ground and

in all

sizes,

ball-type

7-38

not

a common

is accomplished riding

on

valve

ball.

and

the

where

the

design

ball

valves, Many

both oxidizer be sized either

according

to the

flows, The

by lip-

or be

in a

In our specific

case,

7-38

is

which

could

pressure.

actuator

the

balls valves

of a ballvalve

tive

angular

seals,

and

the

valve

seal

Typical open

position,

valves

can

devices,

such

characteristics

of a typical

ball

are

presented

ball

valves

can

by either

The

crank First,

mediate

position

fuel

motion motion

link

arm.

The

actuator

the

valves

are

the

of the the

valve

shown opened

opening),

of

and crank

sequence

by varying

between

(partial

pressure

to a rotary

opening

be adjusted

in

actuator,

reciprocating

of a connecting

positions

stages.

shown

by a piston-type The

by means

arrangement.

the

fully

is translated

two

of the

of the

be powered

gas

arm

surface

90 ° from

K for ball-type

positions

provided

to

or O-ring-type sealing

specific

surface.

rotates

geo-

7-39. activation

figure

of the

The

sealing

Flow

drops

angular

using the

flow-regulating

pressure

affected

through

Ball-type

throttling.

at various

or inert

and fuel individ-

as

position.

at the

as

bellows.

bellows.

ball

open

7-4.

to

an unbalanced

coefficients

determined

for constant

The

a

valve

fully

in table

in figure

mechani-

designed

sealing

to the

designed

bearings.

spherical

the

the

mounted

passes

of the

for propellant

valve

is trunnion

rotation

be used

the

attached

within

surface,

within

resistance as

readily as

of such

are

valve

an

of the

creates

on antifriction

designs,

listed

and

, is

area

sealing

ball

center

closed

(4ds2)

this

axles

of valve

are

arrangement,

elements

size. spherical

How-

designed

valve

metric

ring

enclosed

area

effective

on the

integral

axis

of a seal

closing, pressure

valves, to meet

the

fluid

fluid-flow

actuator.

valve

propellant

are

dual-valve

propellant

The diameter,

acting

by the

the

diam-

difficult

the

force

In most

to

frequently.

valves

cham-

infrequently,

critical,

quite

controlling valves can

shown,

volumetric

larger

applications

used

the

For

valve

Each

com-

requirements.

illustrates

Here,

cally linked, flows. The

have

are

are

respect

than

During

two

Up to a nom-

valves

only

envelope

by a single

as

used

main

seal

be less

storable gas

with

weight.

linked,

Figure

thrust).

increasingly

propellant

valve.

It has

some

thrust

ball

service

are

struc-

service. and

types

are

size

unre-

enhances

for lower

valve

and

mechanically operated

in-line,

also

50000-1b

and

flight

ually

as

valves

is its

for high-capacity

it becomes

weight

use

of 3 inches,

to other

because

it permits

in cryogenic

to about

parable

valve

for high-pressure

a reliable

bellows.

effective

of a ball

low-pressure

ball-type design.

relaaxles

has

two

to an interthen

to the

299

DESIGN OF CONTROLS AND VALVES

*,_l_

vAtvl

oo_Lt7

t

tooor_

I

i

700_

[

,2"-

'

_ svtL

P' I - X-_-__

.._--n-Jh_'_i_-_,,--_

' _-\1"'×

....

400---

=J 2_

30"

.t_"

._"

60"

ANGULAR

70"

gO"

=

_+

POSITION

OPEN

NOTE:

Zero _P2

Figure

f_

from

a5

-'%" PI

0"-

S'due

7-39.-Flow

pressure

lea1

Figure

angular

positions.

for

open

position

using

each

stage.

The

by venting the

tween

valve

are

ports

Dual

pistons

of the

opening

port.

them

sealing

closing

both

closing

separate

at

in the

balance

result varies

in the proper as a function

and

seals

provided

with

for all

for

is

This

type

pressurizing

flow

attd

a drain

cryogenic,

is effected be-

dynamic

Propellant

Figure

7-30

shows

valve.

valve

to the

is

to

sequence main

suitable

storable

propellant

as

which force at

A small

attached

high-pressure,

valve.

for high-

as

well

as

services•

Propellant

Figure

a typical seats.

closed. use

All

of elastomers. of all suitable

of poppet

plicity.

This

operation

of an

ment

requires

sage.

and

permits

actuator. turning

without

throat

of the

nonwiping

this

with

fluorine

their

and

relative

to the

the

direct, flow

Typical

design

A main

results

drops.

coefficients

achieved

due

of the

to use

de-

that

However,

consequently

high-pressure in table

is

is and

propellants. valves

lations,

seals,

for use

is largely

which

nection

dynamic

reactive

vantage

are

valve

nor-

Because

is particularly highly

valve

operated

sealings

characteristics other

poppet

This

to be pneumatically

mally

7-41

in-line

con-

inside

the

pas-

are

given

size

presents

the

(6 to 10 inches propellant hydrodynamic

size

actuator,

The

effective

nominal

valve.

unbalanced of the

design To

forces,

a balance area

and

of a typical,

and

the

thus

flow only.

the

pressure

(say rate

cavitating

to operate

as

Bernoulli's

minimum its

region the

pressure

of a

pressure.

forms

at the

throat.

throat

at the

veloc-

variations

beyond

the

pressure

differentials

across

of upstream

pressure),

venturi

as

on upstream a throttling

affords

smaller

and

throat.

_0 percent

used

a

energy

vapor

pressure

advance

is dependent

When

may

measurements. K for venturi

on

through

cannot

ducting venturi

7-4.

below

downstream

The

pressure

the

designed

the

to fall

moves

of sound,

that

Based

a gaseous

gas

venturi

pro-

is

fluid

a result,

disturbances

the

chamber

be

(7-2)),

made

Up to minimum

diameter), reduce

(eq. is

If this ity

7-40

poppet-type

As

may

venturi.

than in the

Adjacent

for flow

in table

venturi

equation liquid

resist-

valves

The

psi.

coefficients

given

cavitating

smaller

limit

conceivable

resistance are

instalreasons,

solution.

venturi

be used

for a

installed

is a possible

to a few it is

size

A valve

of the

penalties

permitting,

valves

arrange-

fluid-flow

contours

Typical

in relatively

K for poppet

of a venturi

sim-

this

duct.

design

In certain for various

of a nominal

main

simultaneously

reciprocating

a typical valve.

be desirable,

a valve

of the

drop

presents

it may

smooth

Valves

propellant

ad-

7-4.

Figure

vided.

so

seals

Valves

metal-to-metal

signed

large

of the

of poppet

venturi-type

ance

dimensioned

to ambient.

Poppet-Type

the

are

counteracting force of the unbalanced

mechanically

Venturi-Type

with

design.

chamber

positions

valve

valve

poppet-type

valve

constant

ball-type

various fully

large-size

propellant

ol a typical

......

7-40.-Typical

deslgn

characteristics

drops

various

to

\!;

the

pressure device, pressure

the drops,

3O0

DESIGN OF LIQUID PROPELLANT

ROCKET ENGINES

loaded -- VALVE

poppet

line

pressure

is used

,._j_

LET

in the

interconnecting

sure tNLET _---SOLENOID

PILOT

VALVE

same

SOLENO1D

PILOT

VALVE

to open,

valve

poppet

it can

be replaced

_

--------=_

FLOW

main

valve

I CLEARANCE

Figure

_---'-

\

"----VALVE POPP_

AREA

LVALV

VALVE

SCHEMATIC

?-41.-Typical

valve

designed

venturi-type and

Development

are

propellant

manufactured

by Fox

Gate-Type

Valve

Co.

parts

Figure pellant

7-42 gate

stricted since

the

effect

a near

rate

gaseous

and

linear

supply

to a square have

characteristics

been

pressure,

law.

The

venturi

systems

well

valve both,

as with

drop

flow

according

venturi

valves

in cryogenic

which

a shutoff

require

control

a cavitating

at a weight

pressure

than

It also

between tion

of the

7-42

uses

in the

the

flow

valve,

diffuser

of only

of only

the

valves

direction

line

systems. inches

of flow,

the

valve

size

and

However, nominal

built

vehicle

7-41

presents

propellant

valve

Fox

Development

Valve

shutoff divergent

valve

between This

The

is unre-

in low-pressure

a relatively and

outlet

design

short

distance

in the

direc-

shown

O-rings

limi-

will

pro-

and

at a

in figure

as the

valve-seat

ACTLIATrP,_G F'LUI_ INLET

f

PISTO_

TYPE

ACTUATOR

space

been

f

ACTUATOrl

_OD

DYP4,_,I_C O-PiNG

of

imposes

in engine up to

PORT

(O_ENING)

ST_TtC O.R VALVE SEAT

G_,TE t _

_SEAT

SPRING

10 IhlLE T

successfully

/_u=

_

OCTLET

systems. a typical and Co.

consists

section

long

length

valves have

designed

which

venturi

resulting inlet

of a pro-

the

4 to 6 times

application

venturi

diameter

for rocket Figure

a relatively

diameter.

on

design advantage

venturi.

occupy

nominal

limitations

a typical

provides

elastomer

of

reliability.

Its major

flow,

flow.

exist

number

Valves

shows

valve

small

and

INLET

Venturi

the

valvebody

rod and

services.

In fluid-flow as

rather fluid-flow

drop.

throat

between

successfully

propellant

tation

vide

relationship

applied

storable

at the

the

the

valve

no pathways

enhances

valve.

fluid

pres-

venturi

Since

The

Propellant

than poppet

causes

or actuator

seals,

further

the

and

The

seals.

to ambient.

the rate

in a reduced

to open.

no dynamic

is ener-

behind

spring

by a shaft

to as-

valve

through

results

and poppet

at a greater

valve

no dynamic pierced

moving

E SPRING

This the

poppet

for leakage

VENTUR)

out

and

upstream

spring

pilot

by leakage

area.

in a pas-

cavity

valve

the

As the

on the

pressure

is vented

overcoming

there _

the

When

valve

cavity

poppet force

propellant

sure

inserted poppet

the

as

7-41),

Normally,

seating

closure.

gized

is not

fills

direction

valve

is

the

pilot

or closed.

(fig.

throat.

pressure

contains

\

Propellant

open

valve

in venturi

clearance .,,_F---

pilot

additional

in the

valve

schematic

closed

provides

SPRING

valve

sageway propellant

I

throat.

by a solenoid

the

normally

an opening

C/_VITY

at the

controlled

to actuate

shown

POPPET

seated

POPPET

venturi-type

manufactured

by

_" SEAL _R_

\L

0 - RINI_ RET_INEA VALVE GATE GUIDE PLATE

It is a pilot-operated of a convergent-

with

V_LVE G_TIE GUIOE PLATE

a simple,

spring-

Figure

7-42.-Typical

gate-type design.

propellant

valve

301

DESIGN OF CONTROLS AND VALVES

seal.

These

are

temperature other

seal

types

designed psi.

suitable

services. are

valves

are

cations

Gate pressures

limited

to low

as

ground-support

relative

valves

as

generator

appli-

engine

and

into

Needle-Type

Propellant

A typical shown

for extremely

low

a dual-valve

linked

by

a mechanical

an integral

part

assembly.

Both

actuation

type

chambers.

The

yoke.

of the

thrust

valves

are

The

trol

systems,

servo

for

valve

and

body

is

by a quick-response

Sealing

at the

at the

metallic

The

pintle

cating

as

vanes

motion

propellant

rods

is seal

well

as

as flow

of

valve

for the

(2)

No leakage

other

needed

CONTROL

special

for very

PILOT

low

ble

or con-

Impor-

for pilot

valves

are:

Sufficient

The

output

fluid

through

the

closed

actuating

power

systems

design

with

(4)

output

souree

at the

design

valve

can

of a pilot

compatipoint

be defined

W = Pd_'

as (7-17)

where W = pilot

space

VALVES

in turn is used to control or

A

F

iNLET

.......... .

valve

output

at the

design

point,

in-lb/sec

The

control

fluid

discharge

design point, psig volumetric flow rate

point,

in3/sec

most

frequently

may be classified

used

pressure at the

on-off

according

design

pilot

to their

at

valves

design

con-

into

(I)

Two-way

(2)

Three-way

types

(3)

Four-way

types

U(L

i]___'

?',

,;::"

.>

ENGINES

DESIGN INTEGRATION SYSTEM CALIBRATION

I

700

E

ROCKET

Design Engine

Requirements System

Because

FOR

ENGINE

for the Calibration

of unavoidable

mechanical

of an

toler-

ances, it may be expected that the operating characteristics and performance of the various engine system components will deviate somewhat from their nominal design value. A certain amount of calibration is always required for these components, as well as the engine system as a whole, to attain the desired engine performance characteristics within design specification. Therefore, provisions must be made in component and systems design to permit effective calibration during system integration. The specific impulse Is of an engine system is the ratio of thrust F to propellant weight flow rate &. Thus, any deviations affecting F or & will affect system performance. I s also is a function of propellant mixture ratio. It is desirable, therefore, and beneficial to calibrate an engine system by adjusting its propellant feed system. Prior to complete engine system calibration, the pressure or pressure drop versus flow characteristics of each individual component should be calibrated and evaluated. Hydraulic and pneumatic components, such as pressure and flow regulators, valves, flowraeters, ducts, and lines, can all be readily calibrated on flow benches. However, those components which operate at temperature extremes, such as thrust chamber assemblies, gas generators, and turbopumps, are best calibrated by combining the flow tests with actual hot firings. The characteristic propellant flow curve of an engine system is obtained by summing the pressure or pressure drop versus flow curves of the various components (figs. 10-5 and 10-6). The general design approaches toward calibrating an engine system to attain its design thrust at design mixture ratio are: (1) The design operating point of each component should be kept within the relatively flat region of its pressure or pressure drop versus flow curve. (2) The mechanical tolerances and built-in adjustments of each component should be designed so that the random deviation

391

ENGINE SYSTEMS DESIGN INTEGRATION

_MINIMUM _CURVE REGULATOR

OUTLET

PRESSURE

SYSTEM

_r

OE SIGN

POINT--\

LINE

PRESSURE

CALIBRATION

DROP_._

ORIFICE

--

VALVES

AND

/I

LINES_

//

VS r_.OW RESIST/kNCE

ORIFICE

/i

L_NE

_

_

//

{AT

/

_

ENTRANCE

DESIGN

/-_OESIGN

_

./

PLOW

PRESSURE

TO

VS

FLOW

CURVE

TANK)

CHAMBER

INLET

PRESSURE

CHAMBER OURVE

INJECTOR

pressure

feed

VS

END

FLOW

CURVE

PRESSURE

VS

J

PROPELLANT

Figure

PRESSURE SYSTEM

SYSTEM DESIGN PRESSURE VS FLOW CURVE OR SYSTEM RESISTANCE CURVE WITHOUT

_

/

/f

_

__--*_---'-_PRESSURANT

_

DESIGN PRESSURE DROP !N THE CHAMBER ( MANIFOLD COOLING_ PASSAGE AND INJECTOR -_

DESIGN CHAMBER _N,JECTOR END PRESSURE

/ _

_=/-_/

DESIGN PRESSURE DROP-_._ DESIGN PRESSURE DROPS iN

/ /

_'x= _

"_1

REQUIRED TANK CALIBRATED

/ \

PRESSURANT

OR CURVE

_

10-5.-Propellant

design

flow

FLOW,

characteristics

of

(oxidizer v--CALIBRATED \RESISTANCE

FUEL CURVE x--FUEL \CURVE

---....

LB/SEC

a

typical

engine

system

or fuel).

SYSTEM

SYSTEM RE, STANCE WITHOUT ORIFICE

_.

CURVE OXIDIZER

_-

WITHOUT SYSTEM

ORIFICE RESISTANCE

,,, ,---FUEL

PUMP

PRESSUI_:

CALIBRATION ORIFICE

"_,

DESIGN

PRESSURE

_

VS [

DROP

/--FUEL

FLOW

CURVE

PUMP

PRESSURE

AT

NZ

t Pfd

°°° J" OX,Q, i

i FUEL

Wfd

wfo

FLOW

LB/SEC

OXIDIZER

Wfb

%a

FUEL SYSTEM FLOW DESIGN CHARACTERISTICS Figure

IO-E-Propellant

of its

flow

flow

design

characteristics

sign

value

will

able

limit,

in order

calibration, components

from

be kept

and

its

to facilitate

systems

other

design

OXIDIZER SYSTE},IFLOW

de-

a reason-

ratio, ber

design

system

operating

Sufficient

pressure

aside

in each

system due

should

The be

or other

be

propellant

to compensate

to component

tions. then

head

engine

set feed

resistance

design

testing.

Certain

newly

designed

drops

allowed

The

propellant

feed

system

calibrated

by means

adjusting

means.

either

devia-

the

can

and

system

The the lated

for Calibration first

design from

design flow rated

rate

of a Pressure step

is

of each

systems

the

System

determination

propellant,

thrust,

Feed

design

as

of calcu-

mixture

.....

tank

the

versus

chamrates,

components be

may

the

at the

estimated

obtained

specific

from from

have

actual

to be

design

versus

flow

system

of design

is introduced

flow

pressure

system.

flow

for calibration. _ressure

can

pressure

shown

thrust

these

or as

for the by

component as

by actual on

components

summation

an orifice Design

point

design

teristics,

feed engine system.

various

data,

propellant

flow

of orifices

of the

operating

LB/SEC

DESIGN CHARACTERISTICS

verified Based

drops

previous

for contingencies

flow

Is (as

firings).

pressure

region. (3)

and test

FLOW

Wod

characteristics of the A-I stage turbopump

within

to keep

in their

ZER PUMP PRESSURE

can

chamber

design

10-5.

in each The flow

drop

by

versus charac-

In addition,

propellant

minimum

curve

of

pressure

pressure

in figure

curve

be obtained

flow

required

for each

propel-

is thus derived. In most pressure feed systems, the design orifice pressure drop for lant

systems

calibration

_ ----_:::.:=?:

determines

the

maximum

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

392

allowable cumulative pressure drop increase the components above their nominal values. suitable tank pressurization system can then designed, compatible with minimum required pressure versus flow characteristics.

of A be tank

Sam ple Calcu la tion (I0-I) The following data are available from analyses and component tests for the A-4 stage propulsion system, at rated thrust conditions: Thrust chamber injector end pressure range required to maintain rated thrust = 110 ¢ 3 psia Thrust chamber injector pressure drop range (both oxidizer and fuel) = 25 ¢ 2 psi Thrust chamber oxidizer dome pressure drop =3_ + 1 psi Oxidizer line pressure drop = 5-+ 1 psi Main oxidizer valve pressure drop (at the fully open position) = 4 ¢ 1 psi Thrust chamber fuel manifold pressure drop =4_+1 psi Fuel line pressure drop = 4 ¢ 1 psi Main fuel valve pressure drop (at the fully open position) = 4 ¢ 1 psi Pressure allowance required for mixture ratio control by oxidizer valve vernier positioning (fig. 7-4)=¢ 10 psi Determine the design pressure drops of the calibration orifices, and the minimum required tank pressures for design flow rates. Solution The design pressure drop of a calibration orifice must be equal to the sum of the maximum pressure drop increases of components above their design values. Thus: The design pressure drop of the oxidizer calibration orifice = 3 + 2 + 1 + 1 + 1 = 8 psi. The minimum required oxidizer tank pressure at the design flow rate = 110 + 25 + 3 + 5 + 4 + 8 + 10 = 165 psia. The design pressure drop of fuel calibration orifice=3+2+l + 1 +1 =8 psi. The minimum required fuel tank pressure at the design flow rate = 110+ 25 + 4 + 4 + 4 + 8 = 155 psia.

Design System

for Calibration

of a Turbopump

Feed

The propellant flow characteristics downstream of the pump discharges of a turbopump

feed system are similar to those of a pressure feed system. However, the difference in turbopump pressure or head versus flow characteristics from those of a pressurized system dictates a somewhat different approach to systems calibration. For mechanically coupled turbopump feed systems, such as the A-1 stage engine, systems calibration generally involves adjustment of the turbopump speed as well as the installation of an orifice in one of the propellant lines. For turbopump feed systems with dual turbine drive, such as the A-2 stage engine, the calibration can be accomplished by adjusting the speeds of both turbopumps. The design principles for the calibration of mechanically coupled turbopump feed engine systems are best illustrated by a typical example, as shown in figure 10-6. Here, the propellant system resistance curves without orifices (representing conditions downstream of the pump discharges) are constructed based on the designs and test results of the components for the A-1 stage engine system. Next, the discharge pressure versus flow curves of both pumps are constructed from test data obtained with the A-1 stage engine turbopump, operated at speed N1. These pump curves intersect the corresponding system resistance curves at point A. At this speed, fuel flow rate _/fa is above, and oxidizer flow rate _#oa is below the required design flow rates, Wfd and _i,od. To achieve the design oxidizer pump flow _Pod, at a desired discharge pressure Pod, the design operating speed of the turbopump assembly mr, st be raised to a required level N 2 by increasing the turbine gas flow. However, at this speed, the fuel pump, which is mounted on the same shaft as the oxidizer pump, would be delivering a flow rate d'fb considerably above the required design flow rate 1//fd (point B in fig. 10-6). To reduce the fuel flow to _fd, a calibration orifice is placed in the fuel line. This amounts to increasing the fuel pump discharge pressure at constant speed hr2 to Pfc, where _'fd is reached at point C. The pressure drop across the calibrating orifice is represented by Pfc-Pfd, where Pfd is the desired fuel pressure. If fuel flow rate _/fa is below and oxidizer flow rate _/o._ is above the required design flow rates, the calibrating process would be to speed up the turbopump to obtain the desired fuel flow,

393

ENGINE SYSTEMSDESIGN INTEGRATION

and to place an orifice in the oxidizer line. However, it is generally desirable to place the orifice in the system of the propellant with the higher boiling point. In this situation, therefore, and also when the pressure drop across a calibrating orifice tends to become excessive, it is customary to trim the pump impeller so as to reduce the effective speed, and thus attain the required flow and pressure levels. In view of pump efficiency effects, it is desirable to trim the pump drawing the smaller horsepower, usually the one with the lower mass flow rate, except in cases of extreme density differences. The adjustment of the turbine gas flow rate, and thus the turbopump operating speed, can also be made by means of orifices in the turbine inlet gas line, or in the gas generator propellant lines. In general, turbopump feed systems afford less stringent requirements for the various components regarding deviations from their design steady-state flow values, because the system is inherently more flexible. However, systems dynamic characteristics under transient conditions may restrict these deviations. Sample

Calculation

Determine

the location

The following design values and allowable deviations are given for the A-1 stage LOX/RP-1 engine system components, at rated thrust: Thrust chamber injector end pressure= 1095 _+30 psia Thrust chamber injector pressure drop (both oxidizer and fuel) = 200 ± 20 psi Thrust chamber oxidizer dome pressure drop =150+10 psi Oxidizer line pressure drop = 25 -+2 psi Main oxidizer valve pressure drop = 35 ± 3 psi Oxidizer pump specific speed, Ns = 1980 rpm Oxidizer pump suction pressure = 55 psia rain Oxidizer pump discharge pressure at 7000 rpm and a design flow rate of 1971 lb/sec = 1505 -+25 psia Thrust chamber fuel jacket and manifold pressure drop = 290 _+20 psi Fuel line pressure drop = 10 + 2 psi Main fuel valve pressure drop = 15 +-2 psi Fuel pump specific speed, Ns = 1090 rpm Fuel pump suction pressure = 45 psia rain Fuel pump discharge pressure at 7000 rpmand a design flow rate of 892 lb/sec = 1720-+ 25 psi

......LL

drop,

orifice,

and its

ex-

Solution(see sample calculation(6-2)) The required oxidizer pressure head at the design point = 1095 + 200+ 150 + 25 + 35 = 1505 psia. The required fuel pressure head at the design point = 1095 + 200 + 290 + 10 + 15 = 1610 psia. Since the LOX pump discharge pressure is 1505 psia, but the fuel pump discharge pressure is 1720 psia, the calibration orifice must be located in the fuel system. The nominal orifice design pressure drop = 1720- 1610= 110 psi. From a detail analysis, we have found that the change of the fuel pump discharge pressure, as a function of turbopump speed increase or decrease, is a fraction of that of the oxidizer pump discharge pressure. Due to the effects of chamber pressure deviations, therefore, the maximum value of fuel calibration orifice pressure drop is required exist:

(I0-2)

of the calibration

its nominal design pressure pected range of adjustment.

when the following

conditions

(a._) Thrust chamber injector end pressure its lower limit (1065 psia) (_b) All pressure drops in oxidizer are at their higher limits (c_) All pressure their lower

is at

passages

drops in fuel passages limits

are at

(d.._)Oxidizer pump discharge pressure is 25 psi below its nominal value at the turbopump speed commensurate with the stated specific speed (e)

Fuel pump discharge pressure is 25 psi above its nominal value, at the same speed

The equivalent required oxidizer pump discharge pressure under these conditions = 1065 +220+160+27+38+25= 1535 psia. Required oxidizer pump developed head H = 144 × (1535 - 55) _ 2990 ft 71.38 Oxidizer

pump volumetric

Q=

::---

flow rate

1971 ×449 12 420 gpm 71.38 =

394

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

Substitute the required

this into equation pump speed

(6-7) to obtain

N- Ns H°'Ts _ 1980 × (2990) °'Ts = 7190 rpm QOS (12 420) °-s Fuel

pump volumetric

From equation (6-7),the fuel pump nominal developed head at 7190 rpm

I333

or 4730 × 50.45 = 1666 psi 144

flow

892 × 449 = 7950 gpm Q50.45

H = (.NQaS_ \"_-s /

F6970 (7950) °'s11.aaa H = L: _ ] = 4730 ft

= f7170 × _ (7950)°s_ ] 1333 = 4900 ft

The equivalent fuel pump discharge pressure under these conditions would be 1666 + 45- 25 = 1686 psia. The required pressure drop of fuel line calibration orifice under these conditions would be 1686 - 1125- 220- 310- 12- 17=2 psi. Therefore the required range of adjustment for the pressure drop of the fuel line calibration orifice is from 2 to 216 psi.

or 10.4 ENGINE SYSTEM INTEGRATED PERFORMANCE CHARACTERISTICS

4900×50.45 144

-1715

psi

The equivalent fuel pump discharge pressure under these conditions would be 1715 +45 + 25 = 1785 psia. The required pressure drop for the fuel line calibration orifice thus would be 1785-1065-180 - 270 - 8 - 13 = 249 psi. Similarly, a minimum fuel calibration orifice pressure drop is required conditions exist:

when the following

(a_a_)Thrust chamber injector end pressure is at its higher limit (1125 psia), and conditions (b), (c), (d), and (2.) above are reversed The equivalent required oxidizer pump discharge pressure under these conditions = 1125 + 180 + 140 + 23 + 32 - 25 = 1475 psia. Required oxidizer pump developed head H=144(1475-55)=2870 71.38

ft

Substitute this into equation quired pump speed N =1980×(2870)°_ (12420) °.s From equation developed head

(6-7),

(6-7);

=6970

the re-

rpm

the fuel pump nominal

In the process of engine system design integration, an importanttask is the integration of engine system perfonnance characteristics. These data are preparedand compiled by the rocketengine designer to provide the vehicle systems engineer with intbrmationnecessary to integratethe propulsionsystem with the vehicle system. _ere possible,a briefexplanationof the data and itsapplicationshould be included to provide clearerunderstandingand greaterusefulness. The followingare importantaspects of integratedengine performance characteristics.

Nominal Engine Conditions

Performance

Values

at Rated

These are usually prescribed by the engine model specification. These data are for engine system nondnal steady-state operation, at rated conditions. Tables 3-2 to 3-5 are typical examples of nominal engine operating and performance parameters, which include nominal thrust, specific impulse, propellant combination, flow rates, mixture ratio, and various component operating data. Allowable deviations are specified for important parameters such as: thrust, -+3 percent, and mixture ratio, *-2 percent. Engine system specific impulse is usually specified at its minimum value. The performance of all deliverable engine systems must be above this minimum during acceptance tests.

395

ENGINE SYSTEMSDESIGN INTEGRATION

In addition to tables for nominal engine performance parameters, nominal engine performance graphs such as chamber pressure versus engine thrust, and engine specific impulse versus engine thrust, are often included as additional monitoring aid. Figure 10-7 presents a typical performance graph for the A-1 stage engine system, of chamber pressure versus engine thrust at sea level.

Sample

CalcuIation

(10-3)

The following data were obtained from design analyses and component tests of the A-1 stage LOX/RP-1 engine system at nominal rated conditions; i.e., 750000 pounds thrust at sea level: Thrust chamber sea level specific impulse at 1000-psia nozzle stagnation pressure, and a mixture ratio of 2.35 O/F=270 sec Turbine sec Oxidizer Oxidizer cent

exhaust

gas

specific

impulse

Required oxidizer flow for vehicle surization-3 lb/sec Determine the following nominal values at rated conditions: _.) Thrust gas (b) Thrust ber

generated

by the turbine

generated

by the main thrust

_)

system system system

Engine Engine (e_) Engine

-r' u

::)W (_ -r

l

'

A trial-and-error method is used to solve this

] IOO

Engine

oxidizer

mixture

ratio

the corresponding

can now be oxidizer

flow rate 2778x2.35 _i'° = (2.35+ 1) - 1948 lb/sec

Engine

flow rate = 1948 + 3 = 1951 lb/sec

fuel flow rate _i'f= 2778 - 1948 = 830 lb/sec

Oxidizer -

--/--

--

/

I.I000

"

chamber

used to approximate and fuel flow rates:

From equations (6-12) and (6-13), the required oxidizer and fuel pump drive horse power for this approximation are:

1150

1050

propellant flow rates mixture ratio specific impulse

problem. Our firststep is to approximate engine system and gas generatorpropellantflow rates. We substitutethrustchamber Is intoequation (1-28):

The thrust

1200

_'

cham-

Solution

Oxidizer

Z . -- t,t.I n., rr'

exhaust

= F _ 750 000 = 2778 lb/sec Is 270

Fuel pump developed head = 4790 ft Fuel pump overall efficiency =65.9 percent Gas generator O/F mixture ratio = 0.408 Turbine gas available energy content = 359 Btu/lb Turbine overall efficiency = 58.2 percent Required auxiliary drive shaft power = 500 bhp

n,., 0

performance

= 32.6

pump developed head--2930 ft pump overall efficiency = 70.7 per-

1250

tank pres-

--_

CONFIDENCE / _

LIMITS

pump horsepower

1951 × 2930 - 550 × 0.706 = 14720hp

- 830x 4790 = 10 96Q hp 550 x 0.659

[

The corresponding

7OO,OO0 7500(30 8OO,OO0 ENGINE THRUST, POUNDS Figure lO-7.-Chamber thrust at sea level

Fuel

pump horsepower

pressure versus engine [or the A-1 stage engine.

turbine

shaft

horsepower:

Thp = 14 720 + 10 960 + 500 = 26 180 hp From equation (6-19), bine gas flow rate:

the corresponding

tur-

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

396

0.707 x 26180 _i't= 359x0.582

Fuel pump flow rate -88.4

lb/sec =(91.7-

We use this

value

to start

a new calculation

Turbine

cycle to separate main chamber and gas generator data. The thrust generated by a turbine exhaust gas flow rate of 88.4 lb/sec = 88.4 x 32.6 = 2880 lb. Thus the main chamber thrust: Ft = 750 000Thrust

chamber

2880 = 747120

= 14850+

2769 ib/sec

_ 88.4 x 0.408 + 2769 x 2.35 + 3 1+0.408 1+2.35

Turbine

25.6) + (2769-

shaft

Turbine

lb/sec

horsepower

1970.6×2930 550x0.706

Thp-

1942) = 889.8

889.8x4790 + 550x0.659

+500=27090

turbine

(b__)Nominal

main thrust chamber thrust = 750 000 - 3000 = 747 000 lb

(c) Nominal

thrust chamber propellant flow rate = 747 000/270 = 2768 lb/sec

hp

We use this trial:

value

0.707x27090 359 x 0.582 for another

= 91.7 lb/sec

(e_) NominaI

engine

specific impulse = 750000/2860= 262.4

Main chamber

thrust

chamber

Ft = 750 000-2980

= 747 020 lb

flow rate _i'tc =

747 020 27-----0_ = 2768 lb/sec

pump flow rate

_ 91.7 × 0.408 _ 2768 x 2.35 + 3= 1970.6 1 + 0.408 1 + 2.35

Engine Performance Variations Off-Nominal Conditions Engine performance off-nominal conditions

calculation

The thrust generated by a turbine exhaust gas flow rate of 91.7 lb/sec=91.7 x 32.6=2980 lb.

Oxidizer

engine

gas thrust = 92 x 32.6 = 3000 lb

sec

gas flow rate wt-

Thrust

the assumptions

(a__)Nominal

Nominal

exhaust

_ 92 lb/sec

system propellant flow rate = 2768 + 92 = 2860 lb/sec Nominal engine system oxidizer flow rate 92 x 0.408 2768×2.35 + - 1967.7 lb/sec 1+0.408 1+2.35 Nominal engine system fuel flow rate -: (92 - 26.7) + (2768- 1941) = 892.3 lb/sec (d__)Nominal engine system O/F mixture ratio = 1967.7/892.3 = 2.20

lb/sec

Fuel pump flow rate = (88.4-

11 790+ 500= 27 140 hp

gas flow rate w_ = 0.707 ×27140 359 × 0.582

pump flow rate

= 25.6 + 1942+ 3= 1970.6

lb/sec

horsepower

This value closely confirms for the last trial. Thus:

flow rate l_,tc = 747270120

Oxidizer

1941)= 892.1

1970.6x 2930 892.1 x 4790 ThP = 550 x 0.706 + 550 × 0.659 + 500

Turbine

lb

shaft

26.6)+(2768-

lb/sec

Resulting

From

characteristics at various must be available to the

vehicle system engineer. They can be summarized in the graphic form such as figure 2-1 (engine thrust and specific impulse versus altitude curve), or by means of tabulated engine influence coefficients which will be discussed, The effects of off-nominal conditions of the system performance parameters vital for the design of a vehicle (1) Atmospheric pressure (2) Propellant densities (3) Pressures at _he engine (4) Propellant mixture ratio control (5) Vehicle acceleration

-,qmNlm

(6) Throttling

of the engine

following engine are considered system:

propellant inlets and vehicle PU

system

ENGINE

Engine

Influence

SYSTEMS

DESIGN

Coefficients

example,

the total

effects

of several

simultaneously termined

on by

an

summing

the

causes

the

percentage dependent

Dependent

change variables

variables

Engine

thrust,

Engine

specific

Engine Engine

mixture oxidizer

Engine

fuel

of

14.696

and nominal 750000

flow,

892.3

262.4

-0.1780 sec

2.20 ......... 1967.7 lb/sec. lb/sec

psia

for

C 2(po

- pen)+

C 3([

pon Poin)

the

A-1

can

[-

pfn)

Pfn

Cs(Pfi-

Pfin)

+

(10-26) Pfin

where F, F,

: engine system value, lb

Pa,

Pan

Po,

Pon

: atmospheric pressure and its nominal •value, psia : oxidizer density and its nominal value, lb/ft 3

P[,

= fuel density lb/ft s

Pfn

thrust

and its nominal

and its nominal

value,

Poi, Poin :oxidizer pump inlet suction pressure and its nominal value, psia Pfi, Pfin = fuel pump inlet suction pressure and its nominal value, psia C_, C 2, C a, C 4, C s :influence Sample

Calculation

coefficients

(10-4)

Estimate the thrust of the A-1 stage engine system operated at the following conditions, without considering the effects of C* correction: Atmospheric pressure, Pa = 10.2 psia Oxidizer density, Po = 71.00 lb/ft 3 Fuel density, p[= 50.90 lb/ft 3 Oxidizer pump inlet suction pressure, Poi = 65 psia Fuel pump inlet suction pressure, Pfi--49 psia

for

the

C* correction

A-I

versus

variables

Oxidizer density, 71.38 lb/ft

Stage mixture

Fuel a

density, 50.45 Ib,/ft a

Engine ratio and

System curve

shown

nominal

values

Oxidizer pump inlet suction pressure, 56

psia

in fig.

Fuel pump inlet suction pressure,

10-8]

C

_

correction 1.0000

45 psia

values:

lb ..........

impulse, ratio, flow,

Atmospheric pressure,

Pan)+

thrust

C* correction)

Poin

Independent

following

engine

Pan C4(Poi-

For

of

(without

+

de-

from the

of

system as:

Fn

Coefficient

to be obtained

increase variables

be

effects.

lO-1.-Influence

of C* correction

A 1-percent independent

can

change

( F - F n) _ C l(pa-

acting

system

individual

TABLE [Value

influences

engine

the

stage engine be expressed

These are used to convert or correct steadystate, main-stage engine system performance parameters (dependent variables) from one condition to another of parameters (independent variables) such as atmospheric pressure, fuel temperature, oxidizer density, etc. This may be a correction to standard sea-leveI conditions (firststage booster engine), or a conversion to other specified conditions. The coefficients are derived from the linearized solution of a set of steady-state differential equations which describe the performance of an engine system. These equations are solved by a digital computer and presented in tabular form, as shown in table 10-1 for the A-1 stage engine system. Each influence coefficient is expressed as a percentage and represents the change of a dependent engine variable, such as thrust, as produced by a 1-percent change in an independent variable, such as atmospheric pressure. A coefficient preceded by a positive sign (+) indicates that an increase of an independent variable produces an increase in the dependent variable. Conversely, a coefficient with a negative sign indicates a decrease in the dependent variable, as a result of independent variable increase. These influence coefficients are usually sufficiently accurate over the entire design operation range of an engine system. Because the influence coefficients are linear,

397

INTEGRATION

.....

.

-

1.8750

.1780

,2650

.0000 .0000

1.6420 2.0430

.0000

.6530

-0.7420 -

.0640

-1.3650 -1.1120 .3120

0,0440 .0072 .0270 .0465 .0207

-0.0066 - .0150 -

1.1030 1.1350

.0020 .0108

- .0260 - .0632

.0045

.0094

398

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

Solution From equation

(10-26)

(F - Fn)_ (-0.178) F_

and table

10-1, which are for the A-1 engine system. The change of engine mixture ratio is computed for changes in atmospheric pressure, propellant densities, etc., assuming the C* correction first to be zero. For the resultant change in engine mixture ratio, the C* correction is read from the graph. The value of C* correction found is then used with other independent variables to compute the changes in the remaining dependent variables.

10-1:

x (10.2 - 14.696) 14.696

+ 1.875 x (71.0 - 71.38) _ (-0.742) 71.38

x (50.90 - 50.45) 50.45

0.044 × (65- 55) + (-00066) × (49 - 45) 55 45 =0.04531

or 4.531

Engine system Pa = 10.2 psia: F = 0.04531

Sample

percent thrust

of an altitude,

x Fn + Fn = 0.04531

where

x 750 000 + 750 000

: 784 000 lb

Nonlinear

Calculation

(10-5)

Estimate the thrust of the A-1 stage engine system operated at the conditions listed for sample calculation (10-4), adding the effects of C* correction, Also, for the same conditions, estimate the thrust assuming an additional mixture ratio error of +10 percent, due to faulty calibration.

Corrections Solution

When the linear approximation is not sufficiently accurate, the usefulness of the engine influence coefficients can be extended by a technique which allows nonlinear corrections for certain parameters. An example of this method is the C* correction. For instance, a plot of C* correction versus engine mixture ratio change may be used in conjunction with a table of influence coefficients such as figure 10-8 and table

By analogy table 10-1, the change due to tion (10-4) are

with equation (10-26) and using engine system mixture ratio the conditions of sample calculadetermined as

(MR- MRn) _ 1.642 x (71.0MR n 71.38 -_ (-1.365)

(-0.002)

x (50.90 50.45

71.38)

- 50.45)

÷ 0.027 × (65- 55) 55

× (49-45)__0.0162 45

or -1.62

percent

-0,1 _-. Z L_

From figure 10-8, the C* correction for a mixture ratio change of-1.62 percent is approximately -0.02 percent. From table 10-1, the influence coefficient for engine system thrust is 1.1030, for a i percent C* correction. Thus

I

-02

i

Z

9 -o.3 i.-

e¢ _ u

-0.4

(Percent

change

in F) = 4.531 + (-0.02) = 4.509

-0.5

-O.6 -16

-12 CHANGE

-8

-4 IN

ENGINE

8 MIXTURE

12

Therefore, engine system effects of C* correction:

x 1.103

percent

thrust

considering

RATIO,

PERCENT

F= 750000 Figure 10-8.-C* engine mixture engine.

versus change in correction ratio curve for the A-1 stage

× (1 + 0.04509)

= 783 820 lb

If the mixture ratio error of 10 percent is added, the total mixture ratio change = 10-1.62

399

ENGINE SYSTEMSDESIGN iNTEGRATiON

= 8.38 percent. From figure 10-8, the C* correction then is approximately -0.11 percent. Thus (Percent

changein

F)=4.531 =0.441,

Engine

system

+(-0.ii)×

1.103

or 4.41 percent

thrust:

F = 750 000 × (i + 0.0441)

= 783 080 Ib

10.5 MECHANICAL INTEGRATION ENGINE SYSTEMS

OF

Basic Considerations Besides combining allcomponents and subsystems functionallyand physically,the design formechanical integrationof an engine system must considerti_eoverallenvelope of the system and itsweight. This includes the locationof the system's centerof gravity. Also, itshould permit simplifiedmaintenance and checkout practices. Judiciouspackaging design techniques should be applied to minimize the number of interconnectinghydraulic,pneumatic, and electrical lines, with their attendant fittings, connectors, joints, and other potential trouble spots. Welded and brazed joints should be used as much as possible. Problems introduced by vibration, high temperatures and pressures, leakage and space restrictions are thus more easily handled. Engine mechanical integration is a vital part of the system design concept; therefore, all factors related to integration and packaging of components and subsystems must receive careful consideration early in the preliminary design stage. In general, a modular engine packaging approach should be selected such as used for the A-1 (fig. 3-2) and A-2 (fig, 3-4) stage engine systems, as well as for many advanced operational engine systems. This assures engine integrity from time of manufacture through vehicle launch. It also provides a compact package for ease of handling, transportation, and installation in the vehicle. Ease of checkout and component accessibility is also afforded by the packaging concept. The engine should be completely assembled in the manufacturer's plant. Subsequent acceptance testing, air transportation, and installation

in the vehicle in the field should not require assembly of additional major components. Integrity of the propellant feed and hot-gas systems, once verified in a complete system during acceptance test, is not necessarily mdlified by the need to temporarily disassemble the engine for shipment. The integrated engine package concept provides added assurance that static teststand firing results have verified structural soundness of the package to a substantially greater degree, than is the case for a system where the vehicle provides portions of the engine structure. An example of a special case of mechanical integration of a liquid propellant rocket engine is the prepackaged storable liquid rocket propulsion system shown in figure 8-1. This system is a completely integrated assembly of all-welded construction, consisting of thrust chamber assemblies, propellant tanks, pressurization system, and necessary controls. This provides maximum assurance of system integrity from the time of manufacture, which includes loading of the propellants, through delivery, vehicle assembly, and launch. Complete propellant separation until systems start is achieved by hermetically sealed burst diaphragms for maximum safety. Acceptance tests are conducted by taking sample units at random from the production line, and hot firing them. In addition, destructive tests of various types are performed. Packaging

of Rocket

Engine

Components

Most major rocket engine components, such as thrust chamber (fig. 4-1) and turbopump (fig. 6-14) assemblies, readily form a logical, independent mechanical unit by virtue of their function and their physical shape. However, in the case of minor components such as control valves, gas generators and igniters, packaging design principles can best be served by making them an integral part of a major component assembly, or to integrate them by grouping. A typical example is a gas generator assembly externally attached to a turbine inlet flange (fig. 3-2). Similarly, gas generator propellant valves and combustor can be integrated into one unit (fig. 4-51). Certain types of hydraulic and pneumatic rocket engine control components lend themselves most conveniently to the packaging design. Here, one of the main objectives is to

4O0

DESIGNOF LIQUID PROPELLANT ROCKET ENGINES

reduce line runs, by combining all parts and passages into one housing. Such a housing (or mounting plate) is relatively leakproof, trouble areas now being limited to external line connections to other components. Furthermore, if components are packaged in this manner, reductions of weight and size are achieved through the use of common walls and through the elimination of extra mounting platforms, clamps, and fasteners. Since relatively few packages are required as compared to the usually large number of individual components, maintenance of such a system is greatly simplified. Integrated packages are about as easily removed and replaced as are the separate components making up each package. However, the packaged design is not necessarily desirable for every control system. Each case must be carefully studied. As a rule, one or a combination of the following methods is used for packaging engine control corn ponents: (1) Bank packaging: A group of similar flatsided component assemblies are bolted together in a bank or stack, with common porting through the mating surfaces from one unit to the next. (2) Subplate packaging: Attachment of two or more individually housed components to a subplate, so that all ports of the individual component housings lead into the subplate manifold, through their mating surfaces with the subplate, and on to the systems plumbing. (3) Cartridge packaging: Two or more components housed individually in cylindrical cartridges are in turn assembled in a common body with suitable manifolding to the systems plumbing. (4) Multiple-component packaging: Detail parts for two or more components are assembled in a normal fashion in a common housing or body. Figure 10-9 presents a typical pneumatic control package for a large liquid propellant rocket engine. This package combines two pressureregulator assemblies, two relief valves, a series of solenoid valves, filter units, and check valves. It controls the flow of helium gas to various engine components. When engine start is initiated, the helium control solenoid is energized allowing helium to flow through the main pres-

sure regulator to the control system. The helium is routed internally to the main control valves through a fail-safe check valve. This insures that the various engine propellant valves remain pressurized and thus open, should the helium gas supply system fail.

Packaging of Turbopump Feed Engine Systems In earlierhigh-thrustrocket propulsionsystems, some of which may still be in operational use, allmajor engine components were mounted into a cage-shaped thrust mount, which was bolted to the vehicle thrust frame by way of lugs. Figure 2-4 allows several typical examples. With these systems, vehicle steering was accomplished by means of carbon jet vanes protruding into the jet (V-2 and Redstone), or by swiveling the thrust chamber (Thor, Jupiter). In the latter case, the high-pressure feed lines between pumps and injector had to be much more flexible than for misalinements and thermal expansion/ contraction alone. Most advanced liquid rocket engines are tightly packaged. All major components are attached to the main thrust chamber, directly or by means of mounting structures, as shown in figures 3-2, 3-4, and 9-1. Here, the thrust chamber serves as the principal structural member of the entire engine system. For steering, the complete engine package is gimbaled from a gimbal bearing which attaches directly to the thrust chamber dome. The other half of the bearing is attached to the vehicle thrust structure. The low-pressure propellant supply duets must be sufficiently flexible to accommodate the gimbal motions. It is noted that vehicle steering through gimbaling of a single engine or chamber is effective only for the pitch and yaw planes. For roll control, at least two engines are required. For vehicles with a duster of engines, therefore, this poses no difficulties. For single-engine vehicles, special roll-control devices are needed. These may be small auxiliary nozzles, possibly simultaneously used as vernier engines after main-engine cutoff. The use of the turbine exhaust for roll control has also been proposed. Whether the engine attaches to the vehicle thrust structure by means of a thrust frame or a gimbal bearing, either device must be designed to be capable of transmitting the full thrust

401

ENGINE SYSTEMS DESIGN INTEGRATION

5 MICRON FILTERBLEED PRESSURE REGULATOR-_ \ MA,NSTA_F ..........

HIGH

CC1NT _\

fG.G._ BLEED VALVE CONTROL SOLENOID HELIUM CONTROL SOLENOID VALVE / /---'--'--MAIN PRESSURE REGULATOR tlJ_.._/_ ,/._t_

j.---,GN,T,ON

PHASE CONTROL L

_

PRESSURE

VALVE

/

-..-1::;-_:_i'_>_!_

__FAIL-SAFE

__

_

RELIEF VALVE

t

CHECK 10MICRON

_x__LOW

VALVE

FILTER

PRESSURE RELIEF VALVE

HELIUM INLET

PNEUMATIC

VENT

PORT CHECK

CONTROL

PACKAGE

SCHEMATIC

VALVE

PRESSURE RELIEF VALVE

HELIUM IN LET

IGNITION

PHASE

SOLENOID VALVE

CONT.

VENT CHECK

VALVE VENT PORT CHECK VALVE

INSTAGE SOLENOID CONTROL VALVE HELIUM CONTROL SOLENOID VA LVE BLEED PRESSURE REGULATOR

Figure

lO-9.-Typical

G.G

pneumatic

control

package systems.

design

BLEED VALVE

used

in

CONTROL

liquid

SOLENOID

propellant

rocket

VALVE

engine

J

402

DESIGN OF LIQUID

forces

at full

adequate The

gimbal

reserve

bolt-hole

tolerance

pattern

and

must

deviations.

vehicle one

deflection,

for normal

including

an

for side

loads.

permit

adjustment

In general,

attachment

halves

PROPELLANT

engine

must

be

ROCKET ENGINES

for and

designed

for

another. LU|E

Figures

10-10

and

aging

design

tem.

It is a LO2/RP-1,

constant engine

details

chamber

10-11

illustrate

of a typical

pump

fixed-thrust

pressure

package

the

consists

of the

feed

TANK

sys-

engine

control.

OIL

pack-

The

with

basic

following

sub-

packages: (I)

Gimbaled

main

(thrust dizer (2)

elbow,

Turbopump

trol

and

gearbox,

heater, Gas

chamber

injector,

auxiliary

generator

oxi-

mount)

(propellant lube

pumps,

pump,

electric

drive)

assembly

valves,

assembly

dome,

gimbal

assembly

turbine, (3)

thrust

chamber,

regulator

(combustor, and

turbine

coninlet L_N

duct) (4) Main oxidizer duct assembly main oxidizer valve) (5) Main fuel duet assembly

OX_OIZE_ V_LVE

(including Figure

10-I I.-System

(including main

the

packaging

engine

shown

design

in figure

detail

of

10-10.

fuel valve) (6) Turbine exhaust duct assembly

(includ-

ing heat exchanger) (7) Engine

start subsystem

(oxidizer and

fuel tanks, control valves) (8) Turbopump

lube subsystem

Electrical

(11)

Engine

The

(lube oil tank

and fittings) (9) Pneumatic

(10)

control package

system

control thrust

majority

of the

packages

are

major

of the

engine

The

thrust

chamber

frame

through

the

thrust

Mechanical

assembly

component

installed

periphery main

package

frame

thrust

frame

sub-

or at the assembly.

assembly

is attached

a gimbal

Protection

and

within

to

mount.

System

of Engine

Packages It is normal between gine

the

vehicle means

age

in transit

and

shock.

and

contain Figure

lO-lO.-Major

packages rocket

of engine.

component turbopump-[ed

vent

desiccant

holes,

subsystem

warn,

change

liquid

propellant

trusion

of moisture.

require

communication

system

pack-

moisture,

dirt,

applied

propellant and

Certain with

such to valve

inlets,

these

which

of undesired lines,

and

closures

indicators

the

and con-

closures,

plates

of color,

en-

user, must

simple

Frequently, bags

and

engine

cover

elapse

rocket

by the

against

include

and

may

therefore,

the

or storage

openings. through

accepted

to protect

plugs,

years propellant

Design,

These

regulator

to other

and

flight.

sider

caps,

several

a liquid

is completed

its

as

that date

in-

however,

ambient

air

may

ENGINE SYSTEMS DESIGN INTEGRATION

are many

4O3

tasks, such as sequencing

which can

be accomplished

nmch

more effectively electri-

cally than would

be possible by mechanical

means. It is not possible nor necessary,

in the frame-

work of this book, to describe the physical laws

FUEL

and the general fundamentals cuitry. They _CLOSURE

literature. Moreover,

COVER

of electrical cir-

are covered abundantly

in the

in contrast with most other

basic liquid engine subsystems,

the rocket en-

gine designer will try to use commercially able "off the shelf" components system.

However,

avail-

for his electrical

other cognizant

members

of

the design team will have to provide the basic circuit diagram (schematic)

and other data in

support of the installation of the required electrical components. trical system "_'__THRUST

CHA,_SE/_ CLOSURE

last subsystems

EXIT

Characteristically, the elec-

of a rocket engine is one of the to be "frozen" before produc-

tion. This is because

COVE_

sequencing

for start and

stop represents one of the major engine developFigure

10-12.-Various

protective

closure

covers ment activities, often resulting in repeated modi-

for the

engine

shown

in [igure

10-10. fication of the electrical system progresses.

("breathing").

In this

equipped

with

of dry air

only.

desiccant

for installation

Others

will

sible

be left

and

clearly

their

protective

LR79-NA-

the as

left

shows

closures

access

must

be

such

10-12

be

be re-

flexibilityof electrical design. this process analyses

engine

readily with

bright

in place

the

for the

is acces-

recently,

(see sec. 10.2).

location

Electrical

Schematic

The complete

inadof

Rocketdyne

11 engine.

electrical schematic

ground- and vehicle-based sizable drawing.

elements,

fills a

With the aid of figure i0-13,

presents a portion of an earlier engine

discussed

SYSTEM

of a typi-

cal liquid rocket engine system, including its

which

ELECTRICAL

More

has been greatly aided by dynamic

static-firingschematic, 10.6

as development

is therefore placed on the

to vehicle.

until

being

may

to permit

covers

must

marked,

Figure

various

closures

of engine These

to prevent

vertently.

of the

in place

operated.

colors,

the

filters

Some

moved actually

case,

Emphasis

as follows.

the basic features are In ordinary wiring dia-

grams, such as that of a radio receiver, all the All rocket engines electrical system

depend

on some

for their operation.

true for solid systems,

where

type of

This is

at least ignition is

contacts of, for instance, a multiple switch or a tube are drawn

initiated electrically, as well as for liquid sys-

crossovers

tems, in which

overs would

the electrical system assumes

to appear in the same

location, as

they do in reality. This requires numerous in the diagram. become

The

number

wire

of cross-

prohibitive in a typical

numerous additional tasks. As with any common household device, electrical circuits in rocket

engine electrical schematic

engines have caused

troubles, due to poor de-

poses it has long been found preferable to draw

sign, misapplication,

abuse, poor maintenance,

the basic diagram

human

errors, and wear.

Properly applied, how-

and may lead to

confusion and errors. For rocket engine pur-

separately.

so as to show

each

circuit

In this "functional flow diagram,"

ever, electrical circuits can" substantially sim-

the various contacts of a relay, for instance,

plify the operation of a rocket engine, and will

appear in different places, and often away

increase its usefulness

the circuit for the corresponding

and reliability. There

from

relay coil. The

4O4

DESIGN OF LIQUID PROPELLANT ROCKET ENGINES

Figure

I 0-13.-Typical

liquid

rocket

drawing system also will materially simplify later troubleshooting. The diagram, of which figure 10-13 shows a portion, was drawn sequentially from left to right; i.e., circuits which are energized during test preparation and start are shown in the left portion, while those associated with the cutoff sequence appear on the right. In the schematic, connectors are shown as continuous double lines, or portions thereof, running horizontally through the diagram (J16, P16-- receptacle 16 and plug 16, etc.). Each of the contacts is called out by a letter (T, G, K, etc.). All wires are numbered, as indicated. Power buses, like connectors, are shown as horizontal lines, or portions thereof (heavy, single line = positive buses, usually shown near the top of the drawing; and double lines = negative or ground bus). The meaning of the remaining symbols becomes clear by following the circuit at the left of figure 10-13. Plug P5 is shown connected to the main power bus K101 at terminal TB1-8. If certain facility signal contacts are properly closed, such as those verifying "'Cooling water OK," "Firex armed," "Observer on Station," and many more, power returns through plug P5, contact "Z," and is applied to relay coil K31C. The "B" contact of this relay closes a circuit to lamp DS31C, which lights up. The "A" contacts of the same relay are in the chain to the coil of K34C, together with the normally open contacts of several other relays, such as K29C, "Heater Power On," and.

engine

electrical

diagram

(partial).

K28C, "Hypergol Cartridge Installed," as well as the normally closed contacts of cutoff relay KglC. If all contacts are properly closed, the "C" contacts of K34C will cause signal light DS34C to light up. Following selection of the ig-nition power source by means of switch S16C, ignition can now be initiated by means of pushbutton $51C, since the "D" contacts of K34C are now closed, and provided ignition disconnect timer K54C has not picked up (TDPU e time delayed pickup : 0.1 seo). In the diagram, several circuit elements appear which are part of other circuits not discussed. Note that in places two relays are used in parallel (e.g., K16C), if the number of contacts required is too large for one relay. The numbers shown in hexagonal frames refer to the channels of an inking sequence strip chart recorder or equivalent instrument. A special test bus K615 is provided which when energized makes all signal lights go on and thus permits spotting burnt-out bulbs. In earlier engine designs, many of the elements shown in figure 10-13 were installed in an engine-mounted relay box. The trend has been to place as many parts of the electrical system on ground as possible. This is easier with first stages, which start while still connected to ground, or even held down mechanically until released, for brief periods following start, than it is with upper stages which must start and stop, and sometimes restart, some time after takeoff.

Beca_se

of individual

approach

and of

ENGINE SYSTEMS DESIGN INTEGRATION

preferences guards two

for the

required,

designers

different

Specifically

with

and

advised, more

are

The lays,

than

applied

Solid-state

shown

Transistor

as

fulfill

the the

a bias

long

as

in R,,

the

Transistor

is

employed engines.

effecLs.

base

are

in a "turned at a higher

emitter.

This V 2.

is no

appreciable will

off"

be

Thus,

is held

in a "turned

off"

QI

except

that

here

the

by

will

this

turns

until

on,

current

the

At

base

This

(1b,) occurs

max

(10-27)

R 4 is

combined

out

"turns load.

to

potential

voltage,

Ib2 flows

to the

switched

to flow. As Ic base of Q2 is

on,"

This

Vbe2.

of the

base

thereby

occurs

At of

supply-

when

Vo, Rs(Icbol+lcbo2)>(V2+Vbe2)

mode

The

base

the

switch

input

vo

is turned

signal,

the

lock-in

switch

10-15

is immersed module

L

LOAD

The

tiN

shows

shown,

for engine

wiring

metically

potting a small

are

comparable

portion

and

together

connectors,

sealed

box

of an diagram.

with

operation,

the

others

including are

connect-

housed

in a her-

or can.

Although

the

two diagrams

signal

emitted

at pin

the

a module Overall

sequence-controller

modules

required ing

a

solid-state

compound.

after

10-16

engine-mounted

--------_v.

through

of a matchbox.

Figure

"_

INPUT

input

input

feeding

the base of Q, switch.

assembly,

in a potting

static

after by

a typical

Following

dimensions

to those

Ichor//

output

to the

shorting off the

shows

module.

removing

another

accomplished

back

(10-28)

of Q_.

maintain

is

voltage

Figure OUTPUT

off by either

base

(to

is removed)

output

max

Vin , or by using

to ground

signal

METHOD

R2

the

+VI)

resistor

transistor

current

Q2

DIODE

+(

voltage,

on.

Q2 base-emitter

diode. In this case, must be used to turn

IM '4-"

Ix

potential,

is overcome.

into

to turn

it reaches

the

point,

the

i t

flow

the

ing

V2

R4

current

base

(V_),

lowered

Switch

LOCK-IN

applied,

Q,

begins at the

Q2 and

flow

_

L'I

be at ground

base-emitter

transistor

Q,

switch ,

is

the

Again, through

off.

ground, and current Ic, increases, the potential of V 2 and

essentially 2.

fashion,

mode

be seen,

is

essentially

voltage

current

the

When

as

current

Vo-V

point,

flows

Vin

Q,

the

ground.

voltage

bias

volt-

raises

above

remain

raises

emitter.

a bias

current

R3(Icbol)>(Vbel

of this

emitter

will

combined plus

causing

which

Q1 will

This

the

less

voltage

can

the

(Vbel),

leg,

base

input

until

than inserting

V 1 volts

and

flow.

this

emitter

R 2, the

will

by

when

is achieved

As

R I and

potential

no appreciable

of re-

It functions

base

potential

are

much

A circuit

10-14.

held

In

function.

advantage

thus

as

long

When

to

switches

the

and

potential

as

potential,

re-

rocket

in the

emitter

do

continue

voltage,

there

while that at the will remain off.

in a similar

may

of which

a similar

have parts

Q2 is

than

inserting

caution by them-

and

above

in figure

by maintaining potential

arrive

of inter-

applied,

(transistorized)

which

to vibration

type is follows:

number

elements

of today's

moving

sensitive

may diagrams.

to the

types

switches no

V_,

discussed

in several

quiring

age,

engine

to malfunction

solid-state

being

is accomplished

good.

diagram

others,

be at a lower

This

circuit

well-developed

be used

must

used,

circuits

these

subject

harm

safe-

of component

electrical

respect

monitoring since

selves

of interlocks,

type

of a comparable

at substantially

locks

types

and

40,5

are

not

related,

X of receptacle

J-5

in

GROUND

figure Figure

I 0-14.-Typical switch

circuit

rocket or "module"

engine

solid-state

diagram.

start

10-13

would

at contact

Similarly,

the

be suitable

A of plug signal

emitted

P51

to initiate in figure from

pin

engine 10-16.

K of plug

406

DESIGN OF LIQUID

PROPELLANT

ROCKET ENGINES

Electrical

Components

To implement

rocket,

cuitry

such

as

10-13, These

a number are:

Relays

and

Relays

and,

the

engine

one

of components

more

in figure

are

used

recently,

are

command

signals

from

and

to translate

center

cir-

shown

required.

Switches

switches trol

electrical

partially

the

to receive the

solid

usually vehicle them

sequenced actuation signals valves and other elements.

state

low

current

or ground into

con-

properly

to igniters, control In combination with

interlocking relay contacts, bias voltages, valve Figure

lO-15,-Assembled module

soJid-state

be[ore

and

after

position switches, continuity monitors, tempera-

switch

ture sensors, spark plug monitors, voltage sen-

potting.

sors, timers, and other devices, they form an engine-contained P51

in figure

10-16,

with

the

relay,

could

be used

to assure

in the

chain

leading

to the

figure

10-13.

aid

engine

coil

sometimes

of an auxiliary

sponse

readiness

of relay

K34C

in

stop.

*

(

GIN[

Jal

FO_

will execute a

to only two external signals: In practice many

changed

R

logic which

elaborate starting sequence

between

in restart and

more signals are ex-

engine and vehicle and/or

ST*AT

20m$

_l_

P_r

•5,

.

I