SPACE LAUNCH VEHICLES


(Source: ISRO)

High Res GSLV Model Photo (575kb)

High Res GSLV Model Photo (800kb)

GSLV (Geosynchronous Satellite Launch Vehicle) - Mk-I

The GSLV (Geostationary Satellite Launch Vehicle) Mk-I is a heavy communication satellite launcher developed to enable India to launch its own INSAT-class 2,000 to 2,500Kg satellites into Geo-Transfer-Orbit (GTO) for Indian and foreign communication satellite market.

GSLV uses major components that are already proven in the very successful PSLV launchers in the from of the S125/S139 solid booster and the Vikas L40/L35.5 liquid fuel motors. The first development flight of GSLV-Mk1 (GSLV-D1) was successfully launched on 18 April 2001, and the second flight (GSLV-D2) on 8 May 2003. The GSLV-D1 was 49 meter tall, weighing 401 tonne and consisting of three stages. The first stage core motor is same as the one used on PSLV and is amongst the largest solid propellant boosters in the world and carries between 129 to 138 tonnes of Hydroxyl Terminated Poly Butadiene (HTPB) based propellant. It has a diameter of 2.8 m. Its motor case is made of M250 maraging steel. The booster is a five-segment solid rocket motor with HTPB propellant and a composite nozzle. Each segment is 2.8 m in diameter by 3.4 m long. The booster motor fires for 107 seconds and develops a maximum thrust of about 4,628 to 4736 kilo Newton (kN). Pitch, yaw and roll control of the GSLV during the thrust phase of the solid motor is achieved by Single plane Engine Gimbal Control (EGC) ( ±5° ) of the four strap-ons. The GSLV-D1 flight had a backup control from S125’s Secondary Injection Thrust Vector Control System (STIVC) realized by injection of an aqueous solution of strontium perchlorate in the nozzle D2 flight dispensed away STIVC[38]. The injection is stored in two cylindrical aluminum tanks strapped to the solid rocket motor and pressurized with nitrogen.

The GSLV has four liquid propellant strap-on indigenous Vikas motors (L40) based on Ariane Viking-2 engine of SEP, which are ignited on the ground, to augment the first stage thrust. Each of the L40 liquid propellant strap-on motors carries 40 tonne of hypergolic propellant (UDMH and N2O4) stored in two independent tanks of 2.1 meter diameter in tandem, the gas generator fed engine burn for 160 seconds and produces 680 kN thrust. GSLV-D2 onwards use uprated version of the Vikas engine tested in December 2001 that develop higher chamber pressure of 58.5 bar against 52.5 bar in the previous version. This new engine uses UH25 (a mixture of Unsymmetrical Di-methyl Hydrazine and hydrazine hydrate) as fuel and nitrogen tetroxide as oxidizer, its new silica-phenolic throat allows extended duration of burning time. It is estimated to increase the stage ISP by about 7 seconds[39] , raising GSLV's GTO payload capability by 150Kg[40] . The L40H is used as booster strapon motor with 42 tonne fuel and L37.5H used as second stage with 39 tonne fuel.


The S125/S138 solid propellant core stage is ignited 4.6 seconds after confirming the normal operation of each of the L40/L40H stages. The mismatched burn time of the S125 core (100 sec) and the L40 strap-on (160 sec) does not give efficient booster performance (because the expended stage cannot be ejected and the deadweight must be wastefully accelerated) and also requires higher cost of close performance matching of the strap-ons. On the D2 flight this has been narrowed down with longer burn time of larger S139 motor (107 sec) and shorter burn time of the more powerful L40H motor (149 sec).

The vented inter-stage between the first and second stage, enables the firing of the second stage 1.6 seconds before the first stage has completed its thrusting action. This design avoids use of additional systems needed to provide sufficient acceleration between the time before the ignition of second stage takes place and sufficient reduction in velocity of the first stage[41] .

The second stage L37.5/L37.5H employs indigenously built Vikas engine based on the Viking-4A engine of SEP France and carries 37.5/39 tonne of liquid propellant - Unsymmetrical Di-Methyl Hydrazine (UDMH) as fuel and Nitrogen tetroxide (N2O4) as oxidizer, stored in two compartments of a common tank separated by a common bulkhead. The gas generator fed engine burns for 150/136 secs and produce 720/804 kN thrust. Pitch & yaw control is obtained by hydraulically gimbaled engine (±4) and two hot gas reaction control for roll. When the stage is expended retro-rocket provide separation velocity.


The last stage C12 is a Cryogenic Acceleration Block 12KRB[42] procured from Glavkosmos, Russia of 15 tonne mass, 8.7 m long and 2.9 m diameter. Liquid-Hydrogen(LH2) and Liquid-Oxygen(LOX) propellant loading of about 12.5 tonne in separate aluminum tanks, powered by a 73.5-kN KB KhimMash KVD-1 RD-56M cryo engine (with two vernier engines) burns for a duration of about 720 second producing a nominal thrust of 75 kilo Newton, and is restartable. The engine has Indian avionics and software. The cryo engine operates at uprated thrust (9% uprate) mode for initial 60% of the burn period till the payload enters local orbit (i.e. free of gravity), the reset of the time it operates at normal mode transforming the orbit into highly elliptical orbit whose apogee is at 36,000Km while perigee remains at 180 Km.

The 7.8 m long and 3.4 m diameter metallic bulbous heat shield of GSLV, of isogrid construction, protects the spacecraft during the GSLV's passage through the dense atmosphere. It is discarded at an altitude of around 115 Km.

The inertial navigation and guidance system Redundant Strap Down Inertial Navigation System/Inertial Guidance System (RESINS/(IGS) which is housed in the equipment bay computes the inertial position and velocity and guides the vehicle from lift-off to spacecraft injection.

GSLV will be declared operational after one more successful developmental flights (D3). GSLV-D1 successfully launched 1540Kg GSAT-1 satellite into GTO on 18-April-2001 and GSLV-D2 launched 1825Kg GSAT-2 to GTO on 8-May-2003 . Commercial flights C1, C2 & C3 is already budgeted, including long lead-time items for C4, C5 & C6 [43] . Efforts are already on to improve the payload in GTO in progressive steps of 2,200Kg, 2,300 Kg and 2450Kg by 2006 [44]. Other than GTO missions, GSLV can also perform mission to LEO and polar missions.


GSLV (Mk I) Configuration

Launches: 2. Failures: 0. Success Rate: 100% pct. First Launch Date: 18 April 2001, Last Launch Date: 8 May 2003. LEO Payload: 6,200kg. to: 200 km Orbit. at: 19.0 degrees. Payload: 2,250 kg. to a Geosynchronous transfer trajectory. Liftoff Thrust: 795,500 kgf. Total Mass: 414,000 kg. Core Diameter: 2.8 m. Total Length: 49.0 m. Flyaway Unit Cost $: 30.00 million. in 2003 unit dollars.

 

GSLV-D1

  (L40) strapon

GS1 (S125)

GS2 (L40)

GS3 (C12.5)

Payload Faring

Gross_Mass [45]
Fuel_Mass [46]
Empty_Mass
(Stage Mass-Ratio)

45,600 Kg
40,000 Kg
5,600 Kg
(0.877)

157,300Kg
129,000Kg
28,300Kg
(0.820)

43,000 Kg
37,500 Kg
5,500 Kg
(0.872 )

15,000 Kg
12,500 Kg
2,500 Kg[47]
(0.833 )

1,250 Kg

Motor Mass-Ratio

?

?.

?

?

N.A.

Thrust@Vacuum

 

Thrust@Sea_Level

(Burn_Time)

69,388Kgf

 

Kgf

( 160 sec)

479,592Kgf

 

Kgf

(100 sec)

 73,470Kgf

 

N.A

(150 sec)

Normal Mode 7,510Kgf[48]
Upratedmode 8,183Kgf [49]

N.A

(720 sec)(I'st burn 500 sec [50])

N.A.

Specific-Impulse

 

Isp@Vacuum Isp@Sea_Level [51]

 

 

281 sec

248 sec

 

 

269 sec

237 sec

 

 

295 sec

200 sec

  Normal mode: 451.4sec[52]
Uprated mode:
452.3[53]
N.A.

N.A.

Length

Diameter

Dynamic Envelop

 19.7 m

 2.1 m

20.3 m

2.8 m

11.6 m

2.8 m

8.7 m

2.8 m

7.8 m [45]

3.4 m [46]

3.05

Chamber Pressure [56]

Expansion Ratio

 52.5 bar

13.9:1

58.8 bar

8: 1

52.5 bar

31:1

  55.9bar

198 :1

N.A.

Propellant

Chemical

Case material

Liquid

UDMH+N2O4

Aluminum Alloy

Solid

HTPB/AP/Al

M250 Maraging

Liquid

UDMH+N2O4

Steel

Cryogenic

LH2 & LOX

Aluminum Alloy

 

N.A.

Control system

Engine Gimbal Control

Multi-port SITVC

EGC two plane gimballing for pitch & yaw control. Hot gas RCS for roll control.

Two swivellable vernier engine for  thrust phase and cold gas RCS during coast phase.

 

GSLV-D2

  (L40H) strapon GS1 (S139) GS2 (L37.5H) GS3 (C12.5) Payload Faring

Gross_Mass [57]
Fuel_Mass [58]
Empty_Mass
(Stage Mass-Ratio)

45,600 Kg [59]
42,000 Kg
5,600 Kg
(0.882) [60]
157,300Kg
138,000Kg
28,300Kg
(0.820)
44,500 Kg [61]
39,000 Kg
5,500 Kg
(0.876 ) [62]
15,000 Kg
12,600 Kg
2,500 Kg [63]
(0.833 )
1,250 Kg
Motor Mass-Ratio ? ? ? ? N/A

Thrust@Vacuum
Thrust@Sea_Level
(Burn_Time)

78,061Kgf
Kgf
( 149 sec)
483,265Kgf
Kgf
(107 sec)
82,041Kgf
N.A
(136 sec)
7,500Kgf
N.A
(705 sec)
(I'st burn 500 sec [64])
N/A

Specific-Impulse
Isp@Vacuum Isp@Sea_Level [65]

288 sec [66]
255 sec [67]
269 sec
237 sec
302 sec [68]
207 sec [69]
454sec
N.A.
N/A
Length
Diameter
Dynamic Envelop

19.7 m
2.1 m

20.1 m
2.8 m

11.6 m
2.8 m

8.7 m
2.8 m

7.8 m [70]
3.4 m [71]
3.05
Chamber Pressure (Pc) [72]
Expansion Ratio
58.5 bar
?
58.8 bar
8: 1
58.5 bar
?
55.9 bar
198 :1
N.A.
Propellant
Chemical
Case material
Liquid
UH25+N2O4
Aluminum Alloy
Solid
HTPB/AP/Al
M250 Maraging
Liquid
UH25+N2O4
Steel
Cryogenic
LH2 & LOX
Aluminum Alloy

Aluminum Alloy
Number of Engines 4 1 1 1 N/A
Control system Engine Gimbal Control Multi-port SITVC EGC two plane gimballing for pitch & yaw control. Hot gas RCS for roll control. Two swivellable vernier engine for thrust phase and cold gas RCS during coast phase.  

 


GSLV Flights:

GSLV-D1


Flight date & time: 18 April 2001, 15:43 IST, Satish Dhawan Space Center, SHAR, Sriharikota
Payload: GSAT-1 (1,540 Kg)

Flight sequence, result and discussion: Successful GTO launch. Orbit: 181 x 32,051Km. Inclination: 19.2

While the first GSLV developmental test flight is primarily intended for validating the vehicle design and its performance parameters as well as the associated ground infrastructure, the flight opportunity is also made use of to place an experimental satellite GSAT-1 weighing about 1540 kg. GSAT-1 was used to prove new spacecraft elements like 10-Newton bipropellant Reaction Control Thrusters, Fast Recovery Star Sensors and Heat Pipe Radiator Panels to validate them before using them in the ISRO operational ISRO satellites like IRS and INSATs. GSAT-1 also carried two C-band transponders employing 10W Solid State Power Amplifiers (SSPAs), one C-band transponder using 50 W Travelling Wave Tube Amplifier (TWTA) and two S-band transponders using 70W TWTA. The satellite had 119kg (dry mass) unified bipropellant liquid propulsion system made up of a single 440 N Liquid Apogee Motor (LAM) and two redundant networks of 8 22 N RCS thrusters for final orbital placement[73].

GSLV-D1's initial launch on 18 March 2001 was aborted one second before lift-off by the Automatic Launch Processing System when it detected that one of the L40 stage strapon did not develop required thrust, the fault's root cause was "defective plumbing in the oxidizer flow line" that was fixed by replacing it with a spare stage and successful launch 18 days later.

The GTO launch had a launch velocity shortfall of 0.6% (mainly due to Cryogenic stage thrust phase shortfall of 4.1 seconds; 705.8 seconds instead of 709.9) but well within the capability of the satellite to correct the error as it performs orbit-raising maneuver from GTO to SSO whereby the orbital major axis is raised from 18,000Km to 36,000Km. Through a series of six orbital maneuvers conducted between April 19 and 23, the satellite's orbit was raised to near-geosynchronous height with an apogee of 35,665 km, a perigee of 33,806 km and an inclination of 0.997.

During the initial orbit raising maneuver an unexpected problem arose whereby LAM fuel mixture was unbalanced due to a fault in dissimilar fuel tanks used to store the propellant, resulting in wasted fuel as well as spacecraft's center of gravity drifted beyond the main LAM motor gimbal swing limit making the flight uncontrollable. Mission control at MCF failed to notice and stop this serious fault as it unfolded. By the time the situation was correctly assessed the only available way to raise the orbit was to use the LAM in conjunction with 4 smaller RCS (reaction control thrusters) that are less fuel efficient (LAM specific impulse (ISP) of 310 against RCS-thruster ISP of 280), the bold strategy allowed raising the GSAT orbit very close to the intended GSO, in spite of initial fuel loss due to fuel tank flow problem. In the end using the new thruster strategy the spacecraft was short of 10 kg fuel to reach the intended 36,000Km GSO orbit. GSAT is drifting 13.212 deg per day with orbital period 23 hours, apogee 35,665 km, perigee 33,806 km, and inclination 0.99 deg. The fully functional transponders and transmitters on board were deactivated on instructions of the International Telecommunications Union.

G. Madhavan Nair, Director, VSSC, later explained: "Last time we had a problem in the total management of the cryogenic fluid in the upper stage. Some anomaly was observed in terms of fuel consumption and management." The problem was analysed thoroughly by means of a series of tests in Russia and ISRO's laboratories here. "Based on these, we fine-tuned the performance of the upper (cryogenic) stage," he added. [74]

GSLV-D2


Flight date & time: 8 May 2003, 16:58 IST, Satish Dhawan Space Center, SHAR, Sriharikota
Payload: GSAT-2 (1,825 Kg) Cost 150 Cr for launcher and 50Cr fro GSAT-2 [75].

Flight sequence, result and discussion:
Successful GTO launch. Orbit: 180 x 36,000Km. Inclination: 19.2°

The accurate launch injected the satellite at targeted orbit. The autopilot cut off the cryo engine 17 seconds before the last stage fully used the balance 300Kg fuel, indicating true payload capability of at least 2,125 Kg. The increased payload capability as compared to GSLV-D1 was realized by:


1. Improved higher chamber pressure Vikas engine used in booster strap-ons and second stages.
2. Enhanced propellant loading in all stages.
3. Lighter structural elements. (e.g. No STIVC on GS1 core, lighter equipment bay & payload adapter).

This mission flight qualified the high chamber pressure Vikas engines (L40H and L37.5H) with distinctly higher ISP and thrust towards greater payload in future GSLV and PSLV flights.

The GSAT2 satellite scientific & communication payload would greatly benefit fundamental science, novel space applications and qualify components for more sophisticated future satellites.

GSLV-Mk III & IV relative size w.r.t. previous launchers (click on image for Higher Res. version)

GSLV-Mk2

The GSLV Mk-1 launcher uses Russian cryo stage. India paid Glavkomos to develop the technology for cryogenic LOX/LH2 engine that is also the first Russia LOX/LH2 engine. ISRO took up the CUSP (Cryogenic Upper Stage Project) challenge after the United States illegally arm-twisted Russia in April 1992 and July 1993 not to sell the cryogenic technology know-how to India[76]. The U.S. falsly claimed that the sale would violate the Missile Technology Control Regime (MTCR) guidelines since cryogenic technology could be used to propel missiles. Russia, however, agreed to sell seven cryogenic stages and a ground mock-up stage instead of the stipulated five stages and technology.

India didn't seriously embark on the development of either cryogenic or semi-cryogenic engines, though it came very close to sustaining such programs in the early 1970s. For their initial experiments, ISRO scientists worked to build an engine using liquid oxygen (LOX) and kerosene [77]. When India was collaborating closely with SAP in France in the development of a liquid propulsion engine, France appeared to have offered to share its knowledge of HM7 cryogenic engine for a very nominal amount. Again, because of its perceived and overwhelming commitment to the development of VIKAS engine, India appeared to have allowed that offer to lapse. Later ISRO began work on the development of a cryogenic engine in the 1980s when it tested a single element injector generating 60 kg thrust. A one-tonne subscale engine was also realized and tested up to 600 seconds. With this, development of the cryogenic engine for use in the GSLV was initiated in 1994.

GSLV- Mk-2 will be an improved version with greater payload capability (upto 2,250 Kg [78]) on the fourth flight due in 2005 [79]that will use:

  1. Indigenous CUSP (cryogenic upper stage program) with larger fuel capacity (variously reported from 15 tonne[80] to 25 tonne [81 ,82]) and uprated thrust of 9.5 tonne [83] instead of Russian 12KRB's 12.5 tonne propellant and 7.5 tonne thrust. For the 15 tonne stage increased thrust shall be applied during initial 300 seconds after which thrust will revert back to standard 7.5 tonne [84](the last two KRB stages procured from Russia will likely be such 15 tonne stage, that would be approx 1.3m longer). The Indian CUSP reached a milestone when it successfully ran for 1000 seconds [85] on September 14, 2002 at full 9.5 tonne thrust [86], indicating ISRO's CUSP program is geared to realize C20 stage.

The fourth generation INSAT-4 satellites series (7 satellites) has been configured primarily to make use of planned availability of GSLV-Mk II. Due to schedule expediency INSAT-4A will be launched on Ariana.

GSLV-Mk III and Mk IV

The GSLV Mk-III is an entirely new launch vehicle and is not derived from PSLV or GSLV-Mk-I/II. In April-2002, Indian government approved Rs. 2498 crores (US$ 520M) for development of GSLV Mk-III able to launch 4,400 kg satellite to GTO, or 10 tonne to LEO by 2007/2008, with growth potential towards a 6,000kg payload capability through minor improvements.

GSLV Mk-III will be a three-stage launch vehicle with first stage consisting of two S200 Large Solid Booster (LSB) with 200 tonne solid propellant stage, that are strapped to the second stage L110 restartable stage (with 110 tonne liquid propellant & 4-meter diameter). The L110 stage will be first Indian liquid engine cluster design with two [87]improved Vikas engines each of 75 tonne thrust. The improved Vikas engine will use regenerative cooling [88]with superior weight & ISP characteristics. The new S200 booster stage each with 3.4-meter [89]diameter and 25 meter long, would be a scaled up version of mature S125 technology, with estimated enhanced thrust of 785 tonne. The L110 stage will be air lit before the S200 strapon are expended. This would also involve developing a bigger and more powerful C25 cryogenic restartable upper stage with 25 tonne LOX/LH2 propellant, and 20 tonne thrust [90], 4-meter diameter and 8.2 meter long. GSLV-Mk III will have a lift-off weight of about 630 tonne and will be 42.4 meters tall. The large payload fairing of 5-meter diameter and payload volumes of 100 cu meter. Unlike the earlier GSLV types first stage of GSLV Mk-III will not require fins due to availability of adequate control from the large stapon motors.

This new launch vehicle development is a major endeavor for ISRO. Most challenging aspects focus around development of huge S200 Large Solid Booster and the C25 Lox/LH2 cryo stage.

The development work on Mark-III began in October 2002. New facilities will be established at Sriharikota and Mahendragiri to develop the solid boosters, the core liquid stage and the cryogenic stage. A massive plant will come up at Sriharikota to produce solid propellants for Mark III. This will be in addition to the existing Solid Propellant Booster Plant (SPROB) facility at SHAR, one of the biggest plants of its kind in the world. The private and public sector industries taking part in the project too have to augment their facilities for the realization of Mark III hardware [91].

ISRO will establish and commission its new facilities for the project within two years. The first hardware will start rolling out in the second half of 2005, and static and structural tests will begin in 2006. ISRO is aiming the first launch of GSLV Mark III towards the end of 2007 or the first half of 2008 [92].

The subsequent GSLV Mk-IV based on GSLV Mk-III would likely have two additional S200 strapons and bigger 160 to 200 tonne core stage with more Vikas engines in the cluster.

Definition of terms [93]

  1. Thrust(vac - kgf), or Thrust@Vacuum: The thrust the motor generates in a vacuum, expressed in the terms of 'kilograms-thrust (kgf)'. This is not a term officially recognised by the scientifically proper establishment but was used by the Russians in their rocket engineering.
  2. Thrust(sl - kgf) or Thrust@Sea-Level: The thrust the motor would develop at sea level in terms of 'kilograms-thrust'. This may be zero for motors designed for upper-stage operation in a vacuum (see Expansion Ratio, below).
  3. Thrust(vac - kN): For purists, the thrust of the motor in vacuum in officially-correct kiloNewtons (= (Thrust-kgf) * g / 1000) where g = acceleration of gravity on Earth at sea level = 9.80665 m/sec^2.
  4. Thrust(sl - kN): For purists, the thrust of the motor at sea level in kiloNewtons.
  5. Isp-sec, or Isp@Vacuum: The specific impulse of the motor in vacuum. The higher the number, the more efficient the motor. The units here are seconds - because specific impulse represents the kilograms-thrust the motor generates per kilogram of fuel per second of operation (kgf/(kg/sec)) = sec). Important relationships are : effective vacuum exhaust velocity of the motor = Isp * g (this is also the specific impulse in kN thrust terms); fuel consumption of the motor = (Thrust in vacuum) / Isp.
  6. Isp @Sea Level: The specific impulse of the motor at sea level. This is the same fraction of the vacuum specific impulse as the sea level thrust is to vacuum thrust. It can be zero (see Expansion Ratio, below)
  7. Burn time: The total operating time of the motor. For solid motors, this is the more-or-less unstoppable period of thrust until all of the propellant is consumed. For liquid motors, this is the rated thrust duration of the motor for a single operation. This is greater than the thrust time of the motor in actually use on a particular stage. The total tested or rated operating time of a liquid motor between overhauls (if it is reusable) is typically many times the total rated operating time per each use.
  8. Chamber Pressure, Pc-bar: For liquid motors, the pressure of the combustion chamber in bar or atmospheres sea level pressure. The proper definition for the chamber pressure is the throat stagnation pressure (total pressure in the critical cross section). A small loss exists between the injector end of the chamber and the throat. American engineers usually give the throat stagnation pressure, Russians usually gives the injector end pressure. This is usually not stated explicitly and thus may be unknown. For solid motors, the pressure in the motor casing during operation.
  9. Expansion Ratio: The ratio between the area of the combustion chamber exit and the area of the nozzle exit. A large area ratio improves the performance of a motor in a vacuum since the exhaust is expanded further, thus converting potential energy into kinetic energy. However, at sea level a high area ratio can result in flow separation, which can drastically reduce or eliminate the net thrust of the motor.
  10. Gross Mass: Mass of rocket stage at launch. Including interstage with the upper stage or payload.
  11. Fuel Mass: Net fuel in the stage. It does not include the pressurization fluid/gas for pressure fed stage.
  12. Empty Mass or Dead Weight: Essentially Stage-mass less Fuel-mass. This includes dead weight of construction material, avionics, control system, batteries, inter-stage, retro-rockets and fluids used for thrust vector control (if applicable).
  13. Motor Mass-Ratio or Motor Fuel-Fraction: The ratio of the fuel to the total mass of the motor. The most of the non-fuel motor deadweight consist of fuel container and rocket engine, for pressure fed stage the pressurization fluid/gas, container and plumbing also adds to the deadweight.
  14. Stage Mass-Ratio or Motor Fuel-Fraction: The ratio of the fuel to the total mass of the stage. Rocket stage mostly consist of motor, but all component like avionics, control system (including actuators), batteries, retro-rockets, inter-stage(coupling it to upper stage or payload), fluids used for thrust vector control.
  15. Number of Segments: Large solid motors can not be realized in single solid propellant grain, instead they are realized by bolting together single grain segments (the joint filled with passive inhibitor filler), whose interior has been cut to desired overall shape.



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