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RUSSIAN/SOVIET LAUNCH VEHICLE DESIGNERS

K H R U N I C H E V S TAT E R E S E A R C H A N D P R O D U C T I O N S PA C E C E N T E R

15,300

10,400

Payload

∅4,100

Oxygen tank Kerosene tank

50 m

4,110

DM Block

11D58M engine

∅3,700

RD-0214 L I Q U I D P R O P E L L A N T R O C K E T E N G I N E S C H E M AT I C Oxidizer tank (N2O4) Fuel tank (UDMH)

40 m

1 – gas generator valve; 2 – fuel gas generator; 3 – fuel intake valve; 4 – fuel tank pressurant; 5 – starter; 6 – regulator; 7 – valve; 8 – combustion chamber valve; 9 – combustion chamber; 10 – bellows; 11 – stabilizer; 12 – oxidizer tank pressurant; 13 – oxidizer intake valve; 14 – gas generator valve; 15 – oxidizer gas generator; 16 – gas generator valve; 17 – gimbal unit.

30 m

RD-0213 engine 17,050

Second stage

RD-0214 engine, ±45°

Oxidizer tank (N2O4) ∅4,100

Fuel tank (UDMH)

RD-0211 L I Q U I D P R O P E L L A N T R O C K E T E N G I N E S C H E M AT I C 20 m

Engines 1 x RD-0211 3 x RD-0210 4 x 582 KN ±3°15’

1 – gas generator; 2 – oxidizer valve; 3 – pressurant generator; 4 – turbopump assembly; 5 – starter unit; 6 - throttle; 7 – combustion chamber valve; 8 – starter valve; 9 – combustion chamber; 10 – regulator; 11 – gas generator fuel valve

10 m

21,180

First stage

Fuel tank (UDMH) Oxidizer tank (N2O4)

Prior to 1992: RD-253 6x(1,500/1,670 kN) Since 1992: RD-275 6x(1,600/1,750 kN)

Fig. 32. Proton-K Launch Vehicle with DM Upper Stage - Cutaway Drawing 86

RD-253 L I Q U I D P R O P E L L A N T R O C K E T E N G I N E S C H E M AT I C 1 – gas duct; 2 – gas generator; 3, 4, 8, 10, 14 – pyrovalves; 5 – regulator; 6 – turbine; 7 – jet pump; 9, 11, 12 – pumps; 13 – throttle; 15 – nozzle; 16 – combustion chamber

rigidly, others enable longitudinal motion through their sliding joints. The first stage is powered by six RD-253 autonomous cruise engines by NPO Energomash (Chief Designer V. P. Glushko). A hydraulic actuator enables maximum thrust vector deflection to an angle of seven-and-a-half minutes. The deflection is made possible by trunnion attachment of the engine in the vicinity of the chamber throat. The RD-253 engine featuring turbopump fuel supply and oxidizing gas afterburning is fueled by nitrogen tetroxide and unsymmetric dimethyl hydrazine at 2.67 ratio of the components. The engine delivers a ground thrust of 1,500 kN, and a vacuum thrust of 1,670 kN for specific thrust impulse of 2,850 N.s/kg and 3,160 N.s/kg, respectively. With the developed chamber pressure of 15.0 MPa, the engine's burning time in rocket-integrated mode is 130 seconds. The liquid propellant jet engine is ignited by breaking the pyromembranes at engine inlet, following which the components are forced inside the gas generator by the pressurant and hydrostatic fluid column, self-ignite and impose initial rotation on the turbopump assembly (Fig. 32). The second, cylinder-shaped stage includes the adapter, fuel and tail sections. The adapter section is a riveted structure that serves to connect the second and the third stage. The body of the section is made of structural rings, a set of molded profile stringers and the skin. There are four channels in the section's fore end designed to remove gases generated during ignition of the third stage's steering engine. Six powder retrorockets covered by fairings are installed in the rear of the adapter section. The fuel section is designed as a joint block of the propellant and the oxidizer tanks featuring a common intermediate bottom to reduce stage length. The tail section consists of the shell (skirt), the load-bearing cone and the protective shield. The thruster facility of the second stage includes four similar liquid propellant autonomous sustainer jet engines: three RD-0210 and one RD-0211. The engines were designed at the KBKhA Design Bureau for Chemical Automation by S. A. Kosberg's team. Unlike RD-0210, the 0211 engine features tank pressurization units similar to those of the RD-253, specifically, the fuel tank pressurant gas generator unit and the oxidizer tank pressurant mixer unit. All second stage motors are trunnion-fixed in the stage structure, which makes possible electrohydraulic deflection of any to a maximum angle of 3°15'. The second stage engines incorporate the turbopump fuel feed system and provide for oxidizer gas afterburning capability. The propellant components are nitrogen tetroxide and unsymmetric dimethyl hydrazine at a ratio of 2.67. While each engine develops 582 kN vacuum thrust, the specific thrust impulse is 3,201 N.s/kg, the developed chamber pressure is 14.7 MPa, and the burning time is 230 seconds. The engines are kicked-off by a pneumatic starter: the initial rotation is imposed on the turbopump assemblies by the compressed gas stored for the purpose in special

bottles. The engines are cut off by actuating pyro-driven cut-off valves. To ensure fire-and-hold staging, the second stage engines are ignited before the first stage sustainer thrust is cut off: as soon as the second stage thrust goes beyond the residual thrust of the first stage's liquid propellant engines, the pyrobolts that connect the trusswork between stages are initiated to separate the stages, while simultaneously the products of burning exhausted from the second stage combustors act on the thermal protection shield to slow down and push away the burned-out first stage. The rocket's third stage is a cylinder-shaped structure that includes the equipment bay and the fuel and tail sections. The equipment bay is designed as a riveted cylinder made up by structural rings, stringers and skin. The control and targeting system units are attached to the structural rings. Special manholes are provided in the bay to enable access to the equipment. The lentil-shaped oxidizer tank is made up by the middle ring bulkhead and the lower bottom welded to the frame. The oxidizer and the fuel tanks are separated by a common bottom. The fuel tank's lower bottom is cone-shaped and receives the thrust of the liquid propellant sustainers attached to it. The top of the oxidizer tank includes a horizontal damping bulkhead. The oxidizer feed line is placed inside the fuel tank. The tail section is a riveted structure accommodating the four-chamber steering engine. Four powder retroengines are attached to the tail section. The body of the tail section is made up by the skin, two docking structural rings and a set of stringers. In addition, the tail section provides for attachment of the second stage by means of pyrobolts and centering pins. The thruster unit of the RD-0212 third stage block includes the sustainer engine and the four-chamber RD-0214 steering engine. In terms of design and operating principle, the sustainer engine is similar to the RD-0210 engine of the second stage: in fact, the 0212 model is a modification of RD-0210 (Fig. 33). The steering engine (designed with no gas generator afterburning function) includes four chambers, one turbopump assembly, two gas generators and the powder starter. To ensure thrust vector control, the gimbalmounted combustion chambers are deflected by an electric actuator to a maximum angle of 45 degrees. The 30.9 kN RD-0214 steering engine provides a specific thrust impulse of 2,870 N.s/kg by burning the already mentioned fuel components: nitrogen tetroxide and unsymmetric dimethyl hydrazine in 1.8 proportion. The second and the third stages are separated under the thrust load developed by the third stage's steering engine that ignites before the second stage sustainer engines are cut off, in the meantime, six powder retroengines slow down the second stage. The sustainer RD-0213 is cut off at the end of the powered phase, leaving the steering engine to operate alone. That approach enables greater accuracy of the required final velocity of the stage. 87

RUSSIAN/SOVIET LAUNCH VEHICLE DESIGNERS

C O M B AT M I S S I L E S F O R C O N V E R S I O N P R O G R A M S Aerodynamic payload fairing and adapter 34,300

Payload Measurement Instruments Compartment

30 m

28,270

Oxidizer tank

5,713

Fuel tank

Measurement Instruments Compartment

UDMH fuel 14D30 engine, ~20 kN Steering engines

Upper stage

Upper stage fuel tank Adapter section

Mechanisms and Instruments Section N2O4 oxidizer RD-237 Engine

22,337

28,500

N2O4 oxidizer

Second stage

Control system compartment

Payload

Breeze KM upper stage

20 m

Spacecraft

Payload fairing

Payload fairing Payload

1,000

Payload fairing

Steering engines RD-0230 Sustainer engine RD-0229

Adapter section N2O4 oxidizer

Oxidizer tank (N2O4)

Engines RD-0236 UDMH fuel Engine RD-0235 134 KN Oxidizer tank (N2O4)

10 m

Fuel tank (UDMH)

Fuel tank (UDMH)

stage) role of the ascent unit of Strela, thus making any modification of the ground control and targeting system practically unnecessary. The sole new component incorporated into the design of the Strela launch vehicle compared to its prototype is the measuring equipment compartment of the ascent unit. The new compartment sized 2,400 mm in diameter and 550 mm in height accommodates the equipment required to provide telemetry and ground-supported trajectory measurements, the 1st and 2nd stage booster emergency cutoff system, the auxiliary coast flight phase stabilization system and the power supply equipment. The ascent unit that contains the measuring equipment compartment and the payload support pad can be protected optionally by two different fairings: a standard RS-18 nose cone (Ascent Unit 1 option) or a larger cone optimized during RS-18 trial launches (Ascent Unit 2 option). By converting an intercontinental ballistic missile into a space launch vehicle the designers secured high reliability factor of the prototype ICBM known to be have been successfully launched 152 times in as many as 155 launch operations. A joint decision of the Ministry of Defense of the Russian Federation, the Russian Aerospace Agency and the Research and Production Association for Mashine Building designated the new launch site of Svobodny as the main venue for future launches of Strela. In addition, certain launches of Strela can take place at Baikonur with use made of the ground infrastructure already available on site. Based on the existing pattern of Baikonur's ground tracks, payloads can be delivered to 63° inclination orbits. Unlike that, the Svobodny Launch Site provides for satellite insertions into orbits inclined 52°-61° or 90°-97°. Hence, utilization of hardware and processes previously associated with the decommissioned RS-18 ICBM has made possible the development of a light launch vehicle that offers relatively low costs of launching just because newly fabricated nodes or assemblies are very few, and pre-launch processing procedures have been brought to perfection over the years of active defense duty, thus introducing a highly competitive product on the global market of commercial launches.

First stage

DNEPR LAUNCH VEHICLE ∅3,000

∅2,500

Engines 3 x RD-0233, 1 x RD-0234

ROCKOT

Cruise engine RD-264 (4 x RD-263)

STRELA

DNEPR

What is now called Dnepr launch vehicle, was originally the powerful R-36M (15A14) ICBM – the one designated RS-20A under START I Treaty or SS-18 (nicknamed Satan by NATO). The commercial launch vehicle was developed by a team lead by Chief Designer V. F. Utkin at KB Yuzhnoye Design Bureau. With a liftoff mass of 211 tons, the converted product offers high power performance, accuracy of orbital insertion and in-flight reliability. The overall conversion program is being implemented by Kosmotras – an international space company established by KB Yuzhnoye, the Southern Engineering Plant, TsNIIMASH

Central Scientific Research Institute for Machine Building and several other Russian and Ukrainian entities. The configuration of the prototype missile allows development of a launch vehicle that meets every requirement set before a launcher intended to deliver up to 4.0 t payloads into orbit. Dnepr launch vehicle is 34.3 m tall, 3.0 m in diameter. The rocket is fueled by high-boiling propellant components: nitrogen tetroxide and unsymmetric dimethyl hydrazine. The converted launch vehicle includes the first, second and the upper stage, the adapter of the first and the second stages, and the aerodynamic nose cone complete with adapter. All the components of the configuration, except for the nose cone adapter, have been borrowed unchanged from the original RS-20A ICBM. Payload is mounted in the body of the upper stage on a newly designed spacer pad. No modifications were introduced into stage engines. While the 1st stage has four RD-263 single-chamber gimbaled, close-loop motors, the 2nd stage is powered by the RD-0229 single-chamber sustainer and the four-chamber RD-0230 steering motor (block RD-0228). The original high-precision inertial control system operated by the central onboard digital computer has been upgraded to ensure adaptability of the software and compatibility of spacecraft's electrical connections with the ground checking and launching equipment, and to enable input of service pre-launch and in-flight commands. The rocket launches in a mortar mode from its standard launch canister, with the power unit of the 1st stage ignited upon release of the launch vehicle from the canister. «Hot»-mode separation of stages is implemented with the steering engine burning, for which purpose the separated lower stage is slowed down by dropping the pressure of gas in the fuel tank. The payload is separated from the upper stage by burning the upper stage motor in throttled mode for break-off. A feature of Dnepr is the launch vehicle's ability to maintain launch readiness for an unlimited period of time that may be restricted solely by the requirements of the integrated payload. In addition, no supplementary operations are required if launching is rescheduled. So far, mission requirements have not been fully met in only as few as four launches during trial, control or training launch operations in which production Dnepr rockets were involved, bringing the estimated flight reliability factor to 0.97. The commercial operator of Dnepr is Kosmotras Company established in 1997. Dnepr was first launched from Baikonur in April 1999 to deliver UK's UoSat-12 science research experimental satellite to orbit. In its second launch operation in September 2000 five satellites (Italy's, Malaysia's and Saudi Arabia’s) were orbited. The third launch took place in December 2002.

Fig. 38. Rockot, Strela and Dnepr Launch Vehicles: the Products of Conversion 102

103

R U S S I A' S L A U N C H FA C I L I T I E S

L A U N C H FA C I L I T I E S B Y K B O M – T H E G E N E R A L E N G I N E E R I N G D E S I G N B U R E A U

5

4 3

2

6

1 – The N1 (11A52) launch vehicle 2 – Platform 3 – Trunk 4 – Mounting crane 5 – Lightning protection 6 – System for industrial waste collection 7 – Escape slide 8 – Knee 9 – Frame 10 – Idling 11 – Antitheft device 12 – Working stroke 13 – Supporting and rotating gear

7 Top launch umbilical tower (1 item) 8

12

11 10 Supporting truss (ST, 4 items)

1 9

Supporting boom (4 items)

13

Counterweight

Fig. 47. N1-L3 Project Launch Facility

Buffer

Driver of the turn plate

Foundation (4 items) Turn plate

Pilot device (PD, 4 items)

Fig. 46. Baikonur Space Launch Site. R-7 and Soyuz Launch Facilities 126

Lower launch umbilical tower (2 items)

Another element located on the supporting ring is the launch umbilical tower (LUT) designed to bring and connect cables, filling, drain and pneumatic lines or other utilities. Those connections disengage and move away together with the counterweighted LUT during the launch. The rooms available inside the launching plant accommodate stationary propellant filling, thermal conditioning, remote control, compressed gas supply systems, firefighting and gas monitoring equipment, etc. A niche in the launching structure accommodates multiple-level service cabin for servicing the launch vehicle's lower part. The cabin extends above the exhaust duct. The sophisticated equipment of the launching plant is remotely controlled from the command post, where the processes are monitored, documented and displayed by applicable systems. The expertise gained by operating Russia's first space rocket launching plant facilitated designing and construc-

tion of five more plants in 1958-1961, including one deployed at Baikonur and four - at Plesetsk. Long-term operation of the launching plants under climatic extremities of Kazakhstan and Northern Russia, with seasonal variation of the outdoor temperatures from -40°C to +50°C, heavy rain and snow, strong wind proved high reliability of the equipment that survived owing to its robust design and high maintainability. N1, ENERGIA-BURAN LAUNCHING PLANTS. Soviet manned lunar mission program approved in the 1960s required a launch vehicle able to carry at least 100tons payload and a set of relevant ground facilities, including a special launcher. Ground construction works were kicked off in 1964. Within a short time the launching plant featuring unprecedented systems and assemblies was built with its launch system, erector, propellant filling equipment and service tower designed to enable processing, fuelling and launching of the 100-meters tall 127

FOREIGN LAUNCH VEHICLES

US LAUNCH VEHICLES

Payload

3rd stage Centaur IIA Pratt and Whitney RL10A-3-3A engine 134 kN Nose cone Adapter

Designation Boosters 1st stage 2nd stage Length, m 48.77 overall length Diameter, m 3.1 3.0 3.0 Mass, tons 680 tons overall mass Thrust, kN 2 x 6,210 956 463 Specific impulse, N.s/kg 2,602 2,959 3,106 Propellant Solid Aerozine 50/N2O4

T I TA N M I S S I O N PROFILE

4.0

1st stage separation t = 304 s

2 x 67

Liquid Hydrogen

4,350 H2+O2

Liquid Oxygen

N2O4

Orbital flight t = 1,018 s h = 163.2 km v = 7.8 km/s

Control system

Hydrogen valve

LITVC tank

Payload platform 1st stage engine ignited t = 115.9 s

Hydrogen tank Oxygen tank

Helium tank

Hydrazine tank

RL10A-4 PROPELLANT FLOW SCHEMATIC

N2O4

Fairing

Oxygen valve

Boosters separation t = 126 s h = 51.9 km v = 1.87 km/s

Aerozine 50

2nd stage separation t = 541 s

Aerial Fairings jettisoned t = 233 s h = 114 km

Aerojet 1xLR91-AJ-II 463 kN

Guidance system

3rd stage

Pratt and Whitney RL10A liquid propellant engine

CENTAUR UPPER STAGE

Solid propellant engine

Adapter stage

Control system

Aerozine 50

Solid propellant engine Engine nozzle Expanding nozzle

Guidance and Control Guidance: inertial platform and digital computer of General Motors Delco System STAGE 0 1 2 3 Pitch, yaw N2O4 injection Hydraulic Hydraulic By gimballing TVC gimballing gimballing nozzles the 2 nozzles the 2 nozzles Roll idem Hot gas Hot gas Fig. 67. Titan IIIE Launch Vehicle 156

IUS UPPER STAGE

Titan LV liftoff t=0s

Pratt and Whitney RL10 liquid-propellant engine Thrust, kN Mixture Ratio Specific Impulse

RL10A-3-3A 67 5:1 4,444

RL10A-4 85.28 5.5:1 4,490

RL10A-4-1 91.43 5.5:1 4,510

RL10A-4-2 5.5:1 4,510

Aerojet 2xLR-87-AJ-II 956/1,054 kN

Fig. 68. IUS, Centaur Upper Stages 157

FOREIGN LAUNCH VEHICLES

EUROPEAN LAUNCH VEHICLES

Fairing (Oerlikon-Contraves) Diameter: 4 m Mass: 805-915 kg

V=59 m3

Third Stage H10 (Aerospatiale) Mass: 12.5 t Fuel capacity: 10.5 t (O2/H2) Engine: HM-7B, 62 kN thrust

∅3,650

∅3,650

∅3,360 V=28 m3

5,000

∅4,000

2,800

Vehicle Equipment Bay (Matra Marconi Space) Height: 1 m Mass: 520 kg

∅4,000

2,800

6,250

Spelda (British Aerospace) Height: 3.8-4.8 m Mass: 400-450 kg

4,000

V=86 m3

∅4,000

∅2,600

∅4,000

V=49 m3

∅3,650

∅3,360 V=28 m3

N O S E C O N E C O N F I G U R AT I O N O P T I O N S

ARIANE-5 LAUNCHER

2/3 Intermediate Stage UH-25 N2O4

±3°

S E C O N D S TA G E C O N F I G U R AT I O N

∅2,600

1 – helium; 2 – N2O4 oxidizer; 3 – UH-25 fuel; 4 – water; 5 – Viking-4 engine

Second Stage L-33 (Erno, Deutsche Aerospace) Mass: 38,285 kg Fuel capacity: 35 t (UH-25/N2O4) Engine: Viking-4, 786 kN thrust Burn time: 126 s

water

1/2 Intermediate Stage

±3° ∅3,800

∅3,800

First Stage L-220 (Aerospatiale) Fuel capacity: 226 t (UH-25/N2O4 + H2O) Engines: 4 x Viking-5, 4 x 677 kN Burn time: 205 s, ±6°

SSO (Sun Synchronous Orbit) LEO (Low Earth Orbit)

1/2 Intertank Structure

VIKING-5C LIQUID ENGINE P R O P E L L A N T S U P P LY S Y S T E M C O N F I G U R AT I O N 1 – preburner; 2 – turbine; 3 – N2O4 pump; 4 – water feed pump; 5 – regulator

F I R S T S TA G E C O N F I G U R AT I O N 1 – N2O4 oxidizer; 2 – nitrogen; 3 – water; 4 – UH-25 fuel; 5 – 4 x Viking-5C propulsion system

Fig. 89. Ariane-4-4LP Launcher 192

The Ariane-5 heavy launch vehicle was developed by ESA. The development budget is approx. $10 billion. After the first launch of the LV in 1997 failed and the second launch operation in October 1997 resulted in payload delivery into an off-nominal orbit, the third launch a year later (October 1998) was fully successful. The launch operation designated AR504 in December 1999 marked the first commercial event by Arianespace. In total, there were thirteen commercial launches of the basic Ariane-5 version (indexed Ariane-5G-Generic) implemented by Arianespace in 2000-2002. The fourteenth launch performed by an upgraded 10-ton version of Ariane-5 ECA with a cryogenic stage on December 11, 2002 was a failure. The Ariane-5 launch vehicle is able to automatically deliver 18-ton payloads to a low-Earth orbit, or 5.9-6.8-ton payloads – to a geostationary transfer orbit. The Ariane-5 performance for its different main missions is given in the table below. Mission option GTO (Geostationary Transfer Orbit)

PAL Boosters (Erno, Deutsche Aerospace) Mass: 2 x 42.2 t Fuel capacity: 2 x 39 t (UH-25/N2O4) Liquid engine: Viking-6, 2 x 666 kN thrust Burn time: 135-143 s

PAP Boosters (Fiat Avio - Aerospatiale Matra) Mass: 2 x 12.6 t Fuel capacity: 2 x 9.5 t Solid motor: 2 x 650 kN thrust Burn time: 40.5 s Fixed 10°

The fuel charge is a set of four Flexdyne solid-fuel blasters, each one having a star-shaped channel. The solid-fuel booster is separated by means of springloaded pushers with the force of 5.9-6.8 tons. The force provides for the booster to be ejected laterally at a speed of about 5 m/s. The strap-on booster engines are started up simultaneously with the first-stage liquid engine, while the PAP solid propellant booster ignites 3 seconds after the first-stage liquid engine. The commercial operator for Ariane is Arianespace. The launch of an Ariane-4-4L costs between 90-110 million US dollars. On February 15, 2003 Ariane 4 launcher completed its highly successful 15-year career with a perfect mission for its largest customer Intelsat. Demonstrated reliability of Ariane-4 is 97.4 percent with only three unsuccessful launches out of a total of 116 since June 1988.

Orbit parameters Payload Apogee altitude: Single payload: 35,786 km 6,800 kg Perigee altitude: Double payload: 580 km 5,970 kg Altitude: 800 km Inclination: 98.6 ° 10,000 kg Altitude: 500 km Inclination: 28.5°

18,000 kg

The two-stage Ariane-5 has a liftoff weight of around 740 tons. The approximately 52-m high launch vehicle includes two solid propellant boosters strapped on in parallel to the first stage that is powered by a liquid engine. The propellant for the solid propellant boosters (abbreviated as EAP or Etage Acceleration a Poudre) consists of ammonium perchlorate ~ 66 % (the oxidizer) and butadiene (~16 %) and aluminum ~16 % (the fuel). 193

J A PA N ' S L A U N C H V E H I C L E S A N D L A U N C H S I T E S

FOREIGN LAUNCH VEHICLES

payload fairing

40 m payload fairing

payload fairing

third stage: solid rocket second stage: liquid rocket

third stage second stage: liquid rocket

30 m

third stage: solid rocket second stage: liquid rocket LE-5

large payload fairing

second stage: liquid rocket

second stage: solid rocket

20 m

10 m

first stage: liquid rocket

first stage: liquid rocket

3 solid boosters

solid boosters

MB-3

MB-3

first stage: liquid rocket

N-II

90

135

33

35

2.4

2.4

130

350

H-I

Fairing jettisoned t = 226 s

first stage: solid rocket

40 m hydrogen tank

oxygen tank

2 solid boosters 30 m

LE-7

LE-5A engine 122 kN

2xLE-7A

J-I

H-IIA (Augmented)

90

410

33

53

1.8

4

100

4,000

Fig. 94. Japan's Launch Vehicles

multiple advanced design, process or configuration solutions (Fig. 95). H-II is a two-stage launcher with two solid rocket boosters. The second stage incorporates the re-startable main LE-5A engine. While the guidance system does not include a gyro-stabilized platform, the corresponding function is implemented by laser gyroscopes. The first stage of H-II is 35 m long, 4 m in diameter. The mass of the stage is 97 tons, including 85 tons propellant. The propellant components (liquid oxygen and liquid hydrogen) feature the fuel-to-oxidizer mass ratio of 6.0 (Fig. 95). The cruise engine LE-7 of the rocket's first stage with its dry mass of 1,714 kg delivers a thrust of 840 kN on the ground or 1,080 kN in vacuum. The combustor pressure is 13 MPa, the nozzle expansion ratio is 52, while the burning

fairing

liquid rocket booster

2 solid boosters

H-II Liftoff mass, ton 140 260 Overall Rocket Length, m 40 50 Vehicle’s First Stage Diameter, m 2.4 4 GSO Payload Capacity, kg 550 2,200

Payload separation

50 m

time is 346 sec. With the components supplied by means of a turbopump assembly, the turbines of the assembly are driven by the gas produced by the gas generator. The LE-7 engine's pivoting in gimbal suspension enables pitch and yaw control of the vehicle during the flight. Rolling control is effected by two 2-kN thrusters that are fueled by liquid hydrogen borrowed from the cruise engine's supplies. The fuel tank is a thin-wall, cylinder-shape shell with two spherical bottoms. The lower bottom occupies a large volume in the rear compartment. The structure elements of the tank are made of aluminum alloy. The tank covered with thermal insulation coating is pressurized with vaporized hydrogen withdrawn from the combustor cooling line.

oxygen tank

LE-7 ROCKET ENGINE S C H E M AT I C

SRB-A 2x2,300 kN N-I

2nd stage engine ignition t = 362 s

first stage: liquid rocket

first stage: liquid rocket

solid boosters MB-3

LE-5B

third stage: solid rocket

LE-3

1st stage separation t = 356 s

second stage: liquid rocket

payload faring

LE-5A

H-II FLIGHT PROFILE

50,000

large payload fairing

50 m

1 – hydrogen pump; 2 – gas generator; 3 – oxygen pump; 4 – combustor; 5 – nozzle 20 m liquid helium tanks SRBs burn out and separate

hydrogen tank SRB solid rocket boosters: 2 x 1,600 kN 10 m

Liftoff SRB ignition 1st stage engine (LE-7) ignition t=0s

Additional engine hydraulic system tank LE-7 engine 840/1,080 kN

Fig. 95. H-II Launch Vehicle 202

203

FOREIGN LAUNCH VEHICLES

LAUNCH VEHICLES AND LAUNCH SITES OF THE PEOPLE'S REPUBLIC OF CHINA

T H I R D S TA G E 50 m

Fairing

Support Hydrogen

∅4,000 Payload

Oxygen

Avionics Compartment

YF-75 Engine 2 x 78.5 kN

Avionics Compartment

40 m 3rd Stage Fuel Tank

S E C O N D S TA G E

∅3,000 3rd Stage Oxygen Tank

N2O4

Adapter Section

UDMH

YF-75 3rd Stage Engine 30 m

2nd Stage Oxidizer Tank

YF-22 Cruise Engine

Adapter YF-23 Vernier Engine

2nd Stage Fuel Tank Adapter Section YF-23 2nd Stage Vernier Engines

F I R S T S TA G E

YF-22 2nd Stage Cruise Engine 762 kN

N2O4 UDMH

20 m 1st Stage Oxidizer Tank

Apart from the commercial benefits involved in selling launch services and launching international satellites, the People's Republic of China clearly hopes to get access to the most advanced Western technologies. LM-1D LAUNCH VEHICLE. The LM-1D is developed based on the LM-1 launch vehicle that launched the first Chinese satellite in 1970. LM-1D successfully conducted its first flight in November 1997. A three-stage, 31.28 m tall launch vehicle, LM-1D can deliver almost 1,000 kg of payload to a near-Earth orbit. With both the first and the second stages featuring a diameter of 2.25 m, the rocket is topped by a diameter 2.05 m nose cone. The liftoff mass of LM-1D is 85.4 tons. The propulsion system of the rocket's first stage uses unsymmetrical dimethyl hydrazine (UDMH) as the fuel, and nitric acid (HNO3-27s) as the oxidizer. The second stage, powered also by a liquid propulsion unit, is fueled by UDMH and N2O4. LM-1D's third stage is a solid rocket. LM-2C AND LM-2C/SD LAUNCH VEHICLES. The LM-2C launch vehicle is a two-stage, 35 m tall rocket with a liftoff mass of 213 tons. The propellant components are UDMH and N2O4. The diameter of the first, second stage and the nose cone is 3.35 m. The LM-2C first launched in 1975. The LM-2C/SD developed out of the LM-2C differs from the original space rocket by a new upper stage. The rocket successfully launched for the first time in 1997, features liquid first and second stages (UDMH/N2O4) and a solid propellant third stage. LM-2E LAUNCH VEHICLE. A two-stage liquid rocket with four dia. 2.25 m, 15 m tall liquid strap-on boosters. The diameter of the first and the second stages is 3.35 m, the overall liftoff mass is 460 tons, the height is 49.7 m. The LM-2E is topped with diameter 3.8 m nose cone. See Table 19 for more details of the rocket.

Table 19 Parameter Boosters 1st Stage 2nd Stage Design contractor Mass, t Propellant mass, t 4 x 37 Propellant UDMH/N2O4 Propulsion 4 x YF-20 YF-21 YF-22 system (4 x YF-20) YF-23 (vernier) Thrust, kN 4 x 697 Specific thrust impulse, N.s/kg 2,680 2,680 2,911 2,834 (vernier)

LM-3 LAUNCH VEHICLE. In 1984, the LM-3 launch vehicle delivered a 1,400-kg satellite to the 200-35,786 km high, inclination 31.1° orbit. While the lower two stages are fueled by UDMH/N2O4, the rocket's third stage burns H2/O2 components. The propulsion system features the YF-20 liquid engine on the first stage, the YF-22 on the second stage and the YF-73 on the third stage. LM-3A LAUNCH VEHICLE. The three-stage carrier is able to deliver payloads to low-Earth, geostationary or Sun-synchronous orbits. The 52-meters tall, 3.35 m in diameter LM-3A was developed on the basis of the LM-3 using many advanced technologies. Similar to LM-3, the first and the second stages of LM-3A are fueled by UDMH/N2O4, while the third stage propulsion system burns H2/O2. The rocket has a liftoff mass of 241 tons. The first stage incorporates the YF-20B liquid engine delivering a thrust of 4 x 740 kN, the second stage's liquid YF-25 offers 740 kN thrust, the third stage is powered by YF-75 liquid propellant engine for 2 x 78.5 kN thrust. LM-3B LAUNCH VEHICLE. The LM-3B is described as one of China's most powerful launch vehicles with four LM-2E strap-on liquid boosters (Fig. 100).

Four YF-20B Engines

Adapter Adapter

∅3,350

Table 20 China's Liquid Propulsion Systems Booster 10 m

LM-2E BOOSTER N2O4

Oxidizer Tank

UDMH

1st Stage Fuel Tank Fuel Tank

Fairing

YF-20 Liquid Engine 697 kN

Adapter Stage YF-20B Cruise Engines 4 x 740 kN

System Index Propellant Year developed LV (ICBM) Stage Engines on stage Mission control method System configuration

∅2,250 Tail Section

Thrust, kN Specific thrust impulse, N.s/kg Chamber pressure, MPa Burn time, s

YF-3 NT + UDMH 1970 LM-1D 2nd 1 + 1 vernier by steering liquid engine 1 chamber + 1 turbopump 294 (ground)

YF-20 NT + UDMH 1975 LM-2C 1st 4 by chamber swing 1 chamber + 1 turbopump 4 x 697(ground)

YF-22 NT + UDMH 1975 LM-2C 2nd 1 + 1 vernier by steering liquid engine 1 chamber + 1 turbopump 762 (vacuum)

YF-73 O 2 + H2 1984 LM-3 3rd 1 by chamber swing 4 chambers + 1 turbopump 44 (vacuum)

2,620 7.0 135

2,680 8.5 132

2,740 8.5 129

4,169 3.0 451

YF-75 O2 + H2 LM-3B 3rd 2 by chamber swing 2 chambers + 1 turbopump 2 x 78.5 (vacuum) 4,286

Fig. 100. LM-3B Launch Vehicle 210

211

FOREIGN LAUNCH VEHICLES

LAUNCH VEHICLES OF INDIA

INERTIAL NAVIGATION AND GUIDANCE The three-axis attitude stabilization of the vehicle is achieved by autonomous control systems provided in each stage. Single plane Engine Gimbal Control (EGC) of the four strapons of the first stage are used for pitch, yaw and roll control. The core motor of the first stage (S125) is provided with Secondary Injection Thrust Vector Control (SITVC) to augment the pitch and yaw control. The second stage has Engine Gimbal Control (EGC) for pitch and yaw and hot gas Reaction Control System (RCS) for roll. For the third stage, two swivellable auxiliary engines, LH2 and LOX, provided pitch, yaw and roll control during thrusting phase and cold gas system during coast phase. The Inertial Guidance System (IGS) in the Equipment Bay (EB) housed above the third stage guides the vehicle till spacecraft injection. The closed loop guidance scheme resident in the onboard computer ensures the required accuracy in the injection conditions.

Satellite ∅3,400

H2 Fuel Tank ∅2,800

O2 Oxidizer Tank Stage 3

S TA G E 3 ( 1 2 K R B U P P E R S TA G E )

N2O4 Oxidizer Tank Stage 2 UDMH Fuel Tank

Vikas Engine 725 kN

ASLV AUGMENTED SATELLITE LAUNCH VEHICLE The launcher that followed was the ASLV or the Augmented Satellite Launch Vehicle, a rocket by far more powerful than SLV-3. After at least two failure launches in 1987 and 1988, the SROSS-C and SROSS-C2 were orbited by ASLV, respectively, in 1992 and 1994 (Fig. 104). All of ASLV's four stages have solid rocket motors.

PSLV POLAR SATELLITE LAUNCH VEHICLE

N2O4 Oxidizer Tank

∅2,100

UDMH Fuel Tank

Vikas Engines 4 x 662 kN 6°

4,286/4,800 kN

Fig. 106. GSLV Geosynchronous Satellite Launch Vehicle 218

The third series of India's launchers was made by the PSLV (Polar Satellite Launch Vehicle), each around 300 tons in mass, over 40 meters in height. The launchers were able to bring satellites to Sun-synchronous orbits some 800-900 km above the Earth. After the PSLV's maiden launch failed in 1993, the second attempt on October 15, 1994 was a success as the launcher delivered the 300-kg IRS-P2 satellite to orbit. In the third launch in 1996, also successful, another satellite, IRS-P3 of similar mass (300 kg) was injected into the target orbit. The fourth launch of PSLV with more powerful propulsion system took place on September 29, 1997. Similar to earlier launch operations, the launcher lifted off the Sriharikota space launch site some 100 km away from Madras. The latter operation was India's first fully independent launch of the IRS-1D satellite, the fourth in the IRS family after the 1A, 1B and 1C spacecraft orbited under contracts by Russia's launch vehicles (Fig. 105). The successful launch saved India about $20 million. PSLV's payload capacity is growing continuously. In its present configuration dubbed PSLV-C3 the rocket has a liftoff mass of 294 tons, six strap-on boosters, solid first and third stages and liquid second and fourth stages.

Fig. 107. GSLV on Launch Pad

219