complete ppt.ppsx

Submitted by ANSTINE MATHEW AUGUSTINE (32208101006) ARUN KRISHNAN. U (32208101009) MAHESH. J (32208101029) VETRI SELVAN. S (32208101057)

Military

aircraft designed to attack ground and sea target by dropping bombs on them.

Strategic

bombers are designed for longrange bombing missions against strategic targets to damage enemy nations war effort .

Light

bombers Medium bombers Dive bombers Fighters bomber  Ground attack aircraft  Multi role combat  aircraft 

Major type of aircraft designs •

Conceptual design

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Preliminary design

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Detailed design

Conceptual design •

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It depends on what are the major factors for designing the aircraft. (a) Power plant Location: The Power plant must be located in the wings. (b) Selection of Engine: The engine should be selected according to the power  required i.e., thrust required. (c) Wing selection: The selection of wing depends upon the selection of  (1) Low wing (2) Mid wing (3) High wing - For a bomber the wing is mostly high wing configuration and anhedral. - Sweep may be required in order to reduce wave drag.

2. Preliminary design: Preliminary is based upon number of  factors like Loitering.

3. Detailed design: In the detailed design considers each & every  rivets, bolts, paints etc. In this design the connection & allocations are made.

 To

design a bomber aircraft  Range of 20000 km & must carry 75000 kg+ of bombs & missiles.  At supersonic & subsonic regimes  To operate at regional bases with low cost of  operation & maintenance  The aircraft must also be capable of single pilot operation scenario.  Due to long range pilot work load must be reduced  The aircraft must be all weather , all terrain operation capable including the airbase.  To take up a load factor +8g to +7.5g to -3.5g.

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Collect data of existing aircraft of similar purpose i.e., bomber.  The basic factors of aircrafts performance viz. Weight, Cruise velocity ,Range ,Wing area & Engine thrust.  The performance data of various bomber aircraft with payload capacity between 5000 & 56600 kg was collected.

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Mirage IIIE Mirage IVA F-111F F-111F swept  Tu-22R  Tu-85/1  YB-60 B-2A etc

Preferred Configuration:

rom Comparison Parameters

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Max takeoff weight (kg)  Thrust to weight ratio Aspect ratio Wing loading (N/sq.m) Span to height ratio Span to length ratio Combat radius (km) Pay load capacity (kmph) Max Speed (kmph) Service ceiling (m) Max Speed (m/s)

Values

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500000 0.28 8.4 7848 5 1.5 5000 75000 1000 15000 277.77

G enera rrough oug General estimate

Mass Fraction Payload

0.15

Fuel

0.45

Structure

0.32

Power plant

0.07

Fixed equipments

0.01

Total

1.00

Redefined Mass Estimation

h

6’

2

10000 km

9000 km

7’

3

1000 km

1/2 hr 

2’

8’

R  1000 km

3’

0

1

4’

5’

9’

10’

Mission profile for Strategic bombing

Analysis of mission profile TSFC values for Bomber Cruise

Loiter

0.5

0.4

Comparative data of  Engines

Engine Selection Name of the Engine

GP-7000

Manufacturer

Engine Alliance

Type

 Turbofan 2 Shaft

Length (m)

4.74

Diameter (m)

3.16

Wet weight (kg)

6800

Dry Weight (kg)

6712

Maximum Thrust (kN)

363

Overall Pressure Ratio

43.9

Thrust to Weight Ratio

4.73

Fan Diameter (m)

2.95

The above engine has been selected selected from from a list

Redefined Thrust to weight ratio

AIRFOIL SELECTION

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Airfoil nomenclature Lift coefficient Drag coefficient  Types of airfoil Formula used Airfoil

AIRFOIL NOMENCLATURE  The

cross-section shape obtained by the intersection of wing with the perpendicular plane is called airfoil.  The major design feature of an airfoil is the mean chamber line ,which is the locus of points halfway between the upper and lower surface ,as measured perpendicular to mean chamber line itself  .   The most forward and rearward points of the mean chamber line are the leading

 THE

FORWARD AND REARWARD POINTS OF THE MEAN CAMBER LINE ARE THE LEADING AND TRAILING EDGES.

CHORD

LINE THE STRAIGHT LINE CONNECTING THE LEADING & TRAILING EDGES. MEAN

CAMBER LINE THE LINE BETWEEN UPPER &LOWER SURFACES. SU RFACES.

CHAMBER

MAXIMUM DISTANCE BETWEET THE MEAN CAMBER LINE & THE CHORD LINE .

LIFT COEFFICIENT •

 The lift coefficient ( CL or C Z  ) is a dimensionless coefficient that relates the lift generated by an aerodynamic aerodynamic body such as a wing or complete aircraft, the dynamic pressure of the fluid flow around the body, and a reference area associated with the body. It is also used to refer  to the aerodynamic lift characteristics of a 2D airfoil section, whereby the reference "area" is taken as the airfoil chord. It may also be be described as the ratio of lift pressure to dynamic

Drag Co-efficient:  The drag coefficient (commonly denoted as

Cd, Cx or Cw) is a dimensionless quantity  that is used to quantify the drag or  resistance of an object in a fluid environment  such as air or water. It is used in the drag equation, where a lower drag coefficient  indicates the object will have less aerodynamic or hydrodynamic drag. The drag coefficient is always associated with a  particular surface area.

 TYPES OF AIRFOIL •

CHAMBERED AIRFOIL

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SYMMETERICAL AIRFOIL

CHAMBERED AIRFOIL •

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It is also called as unsymmetrical airfoil . Upper surface of the airfoil is not equal to lower surface. SYMMETRICAL AIRFOIL:

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Surface above the chord line and below the chord line are equal.

FORMULA USED

FORMULA USED

6. Airfoil selection and Wing Geometry estimates •

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Main Parameter Selection: Wing Loading:

 Thickness based Reynolds Number

Flap selection:

Wing geometry

Critical Mach number for the airfoil

LANDING GEAR

 TYRE SELECTION •

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Load Distribution  Typical load of aircraft while landing ;WL=W T -O.8WF While aborting mission ; WL=W T-O.1WF During static condition ; WL=W T

CONTACT AREA

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Ww=Ap x P

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Ap=2.3 √ dwww(dw/2-Rt)

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Rt=(dw/2-Ap/(2.3 √dwww))

*RUN WAY LOADING Runway loading=load on each wheel/area of  contact

Runway Loading

DIMENSIONAL ESTIMATES •

Span to height ratio=b/ha ≈5

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Span to length ratio=b/la ≈ 1.5

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CONFIGARATION OF TAIL

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Horizontal stabilizer

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Horizontal stabilizer sizing 15% of wing area;sh/s=0.15 area;sh/s=0.15

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Vertical stabilizer geometry

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Vertical stabilizer sizing 9% of wing area ;sv/s=0.09

Configuration of  tail

Airfoil NACA 0012

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PREPARATION OF LAY OUT Wing location and C.G estimation Wfuselage X fuselage + W wing (X +X wing) = (Wfuselage+Wwing) (X+Xfinal) Where X is the location of wing root L.E from the nose fuselage and Xfinal is the reaction of cg from the L.E at root X final=0.35(Xcr - Xct)

Wing Detail for cg estimation

Three views of  Aircraft

Front view

Side view

DRAG POLAR •

Drag equation for entire Aircraft:Cd=Cdwing+Cdothers+KCL^2

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*wetted surface area

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Fuselage =Wfuselage*hfuselage =Wfuselage*hfuselage

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Engine =4* π/4d^2

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Nose landing gear=dw*Ww*4

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Main landing gear=dw*Ww*12

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Main landing gear=dw*Ww*8 Flap=Lflap * Wflap

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 Take off performance =Cdpermanent+CdLG+Cdflap+Cdwing Landing performance=Cdpermanent+CdLG+Cdflap+Cd wing

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Cruise performance=Cdpermanent+Cdwing

Drag polar

Lift to Drag Ratio

Performance Calculations •

Thrust required and Thrust available analysis: W1= 25% of Fuel and 100 % of Payload

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W1= 3185533.292 N

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W2= 50% of Fuel and 100 % of Payload

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W2= 3784962.23 N

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W3= 75% of Fuel and 100 % of Payload

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W3= 4384391.173 N

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Thrust scenarios at Sea level for different weights

Thrust scenarios at 11 km altitude for different weights

Thrust scenarios at 25 km for different weights