Geosynchronous Satellite Launch Booster
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The glorious journey began at the magnanimous
St.Mary Magdalene Church (left),
it was the first lab and office of ISRO
Currently maintained as a majestic Space Museum - Image: ISROHQ
THE EVOLUTION OF INDIA’S ROCKET PROGRAM
In 1962 in a rather nondescript fishing hamlet called Thumba near Thiruvananathapuram close to the sea, two stalwarts and visionaries of the Indian Space and Nuclear program were scouting for a new rocket launching facility along with other fellow scientists. One was the blue-blooded Dr. Vikram Sarabhai and the other the aristocratic Dr. Homi Jehangir Bhabha. Within a year’s time the Thumba Equatorial Rocket Launching Station (TERLS) was born. This was just a humble beginning but over the years, India achieved the capability of building its own launch vehicles, remote sensing, communication, defense and navigation satellites, launch interplanetary and lunar missions, design and develop the closely guarded and cutting-edge cryogenic engine. TERLS primarily focused on launching sounding rockets (SR) to study of equatorial electro-jet phenomena, which is a stream of electric current flowing in a narrow band on either side of the magnetic equator at a height of around 100 kms.
Contrary to what most of us know and believe the United States was one of the first countries to provide key technologies for our space program such as supplying and training to launch sounding rockets (Nike-Apache) and establishing the TERL station, though years later they scuttled our progress when we decided to develop the GSLV and the cryogenic engine. France also offered invaluable help with their Centaure sounding rockets and they helped India to indigenise it, they also helped ISRO to setup rocket fuel production facilities; Russia and UK also helped us with sounding rockets, radars and other equipment.
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| In the early 70's, due to lack of proper facilities, rocket cones were carried on bicycles. APPLE Satellite being transported in a bullock cart (1981) - Image: ISRO |
INDIA’S FIRST LAUNCH VEHICLE – SLV-3 - INDIA ENTERS THE SPACE AGE
SLV-3 Launch Vehicle & the Rohini Satellite mated to 4th stage - Image: ISRO
As the Indian space program grew in the late 1960s the Indian National Committee for Space Research (INCOSPAR) became the Indian Space Research Organisation (ISRO), and in 1975, ISRO launched its first satellite the Aryabhata, on a Soviet Kosmos B1 rocket.
The transition from building and launching sounding rockets (SR) to a full-fledged satellite launch vehicle is a daunting task. There is a humongous amount of difference between a sounding rocket and a launch vehicle (LV), the SR’s are smaller, lighter and less complex to build and operate. A LV on the other hand, has multiple stages, spent stage jettisoning programs, next stage precision ignition, healthy levels of navigational sensors, control and guidance systems for normal and in-flight course correction, optimized structural build to counter atmospheric flight regimes, aerodynamic abilities, hardware and software cohesion, and most importantly capability to inject a satellite at the right height, angle and speed.
SLV-3 while being a simple launcher was 22 meters high and weighed 17 tonnes, yet it had 44 major systems and sub-systems, innumerable electrical and electronics components, 25 Kms of wiring and 40,000 fasteners. On 18 July 1980 SLV-3(E)-02 the four stage all-solid propellant rocket soared into space and injected the 40 Kg Rohini satellite into a 300 Kms by 900 Kms elliptical orbit thus making India a space power.
ASLV Lift-Off - Image: ISRO
This launcher acted as a low cost intermediate vehicle to demonstrate and validate critical technologies and provided valuable inputs for future programs. With a lift off weight of 40 tonnes, the 23.8 m tall ASLV was a five stage, all-solid propellant vehicle with a mission of orbiting 150 kg class satellites into 400 km circular orbits. The strap-on stage consisted of two identical 1-meter diameter solid propellant motors. (Text: ISRO)
POLAR SATELLITE LAUNCH VEHICLE (PSLV)
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| PSLV on Launch Pad PSLV Stages - Image: ISRO |
The PSLV is the workhorse rocket of ISRO with an outstanding success rate among medium lift rocket systems of the world. The success of the POLAR Satellite Launch Vehicle (PSLV) is an important milestone for the Indian space industry. India has achieved self-sufficiency in launching its operational satellites. The PSLV is a unique vehicle since it employs both liquid and solid fuel engines. The PSLV can place a satellite weighing about three tonnes in low earth orbit (LEO), at a height of 400 to 600 km. It can also deploy satellites weighing up to 1,500 kg in Polar Sun Synchronous orbit at a height of 750 km above the earth, as it has done in the case of the Indian Remote-Sensing Satellites (IRS) constellation. This versatile and flexible vehicle can handle a LEO, a Polar and a GTO orbit.
VARIANTS ILLUSTRATION
DESCRIPTION
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PSLV-CA
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PSLV-G
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PSLV-XL
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Number of Solid
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None
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Six (9t)
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Six (11t)
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Payload to Sun Synchronous Polar Orbit
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1150 Kg
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1600 kg
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1750 kg
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Payload to Geosynchronous Transfer Orbit
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590 Kg
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1000 Kg
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1050 Kg
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VARIANT
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LAUNCHES
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SUCCESSES
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FAILURES
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PARTIAL FAILURES
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PSLV-G (Standard)
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11
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9
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1
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1
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PSLV-CA (Core Alone)
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9
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9
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None
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None
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PSLV-XL (Extended)
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5
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5
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None
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None
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The Geosynchronous Satellite Launch Vehicle (GSLV) is the most important rocket developed by the Indian Space Research Organisation (ISRO). GSLV presented the Indian space program its most exhaustive and demanding technological tests. It embodies decades of struggle in the development of propulsion fuels, engines, avionics, electronics and control systems and cryogenic stage development by teams of dedicated research scientists. India realized the need for a heavy lift booster in the early eighties, as the PSLV was inadequate to place heavy payloads in Geo Synchronous Orbit. The purpose of the program was to satisfy India’s needs for telecommunications, environmental monitoring, disaster warning and other satellite systems as well as facilitating the country's entrance into the world space market by building a massive satellite network and ensuring a credible launch capability. After two developmental flights, the GSLV was declared operational.
GSLV STAGE DETAILS
First Stage (GS1)
The First Stage of the GSLV comprises a solid propellant motor (S139) derived from the PS1 Core Stage of the PSLV and four liquid propellant strap-on motors (L40H). The S139 stage is 20.1 m long and 2.8 m in diameter and it carries 138 tons of Hydroxyl Terminated Poly Butadiene (HTPB) solid propellant. The stage develops about 4736 Kilo Newton (Kn) thrust and burns for 107 seconds. The four strap-on (L40H) stages are 19.70 m long and 2.1 m in diameter and are loaded with 42 tonne of Unsymmetrical Di-Methyl Hydrazine (UDMH) hypergolic fuel and Nitrogen Tetroxide (N2O4) as oxidizer. Each produces 765Kn thrust and burns for 149 seconds.
Second Stage (GS2)
A single Vikas liquid propellant engine powers the second stage (GS2). The second stage is 11.6 m long and 2.8 ms in diameter. It is loaded with 39.3 tonne of UDMH and (N2O4) in two compartments of aluminium alloy tank separated by a thin metal sheet known as common bulkhead. The Vikas Liquid fuel engine produces a thrust of about 804Kn and burns for 136 seconds.
Third Stage (GS3)
The third stage of GSLV uses a Cryogenic Upper Stage (C12). This stage employs liquid hydrogen and liquid oxygen as fuel and oxidizer respectively, it is 8.7 m long and 2.9 m in diameter. Liquid hydrogen (LH) and liquid oxygen (LOX) are stored in two separate aluminium alloy tanks connected by an inter-stage structure. With a propellant loading of 12.6 tonne, the stage can burn for duration of about 705 second producing a nominal thrust of 73.5Kn.
Payload Fairing
The Aluminium Payload Fairing, which is 7.8 m long and 3.4 m
in diameter, protects the vehicle electronics and spacecraft during its
tumultuous ascent through the atmosphere. It is jettisoned when the vehicle
reaches an altitude of 115 Km.
Separation Systems
- GSLV employs various separation systems such as Flexible Linear Shaped Charge (FLSC) for the first stage
- Pyro actuated Collet release mechanism for second stage
- Merman band Bolt Cutter separation mechanism for the third stage
- Spacecraft separation is by Spring Thrusters mounted at the separation interface
GSLV Control Systems
- First Stage: Multi-Port Secondary Injection Thrust Vector Control (SITVC), EGC single plane gimbaling.
- Second Stage: EGC two plane gimbaling for pitch and yaw control. Hot gas reaction control systems for roll control.
- Third Stage: Two swivel able vernier steering engines for thrust phase control and cold gas reaction control systems for coast phase control.
- Inertial Guidance System (IGS): IGS in the Equipment Bay (EB) housed above the third stage guides the vehicle till spacecraft injection.
SALIENT FEATURES
SPECIFICATION
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PARAMETERS
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Height
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49 meters (161 feet)
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Diameter
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2.8 meters
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Lift of Weight
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414 Tonnes
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Strap-on Boosters
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L40H x 4
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Stage 1
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GS1
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Stage 2
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GS2
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Stage 3
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GS3
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Mass to Low Earth Orbit
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5,000 kg
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Mass to Geostationary Transfer Orbit
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2,500 kg
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STAGES AT A GLANCE
PROPULSIVE STAGES AT A GLANCE
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Parameter
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GS Ist Stage ( First Stage )
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GS IInd Stage
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GS IIIrd Stage
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S125 Booster
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L40H Strapon
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Length (m)
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20.1
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19.7
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11.6
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8.7
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Diameter (m)
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2.8
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2.1
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2.8
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2.9
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Total Mass (tons)
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156
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46
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42.8
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15
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Propellant Mass (tons)
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138
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42
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39.3
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12.6
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Case Tank Material
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M250 Steel
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Aluminium Alloy
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Aluminium Alloy
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Aluminium Alloy
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Propellant
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HTPB (1)
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UDMH '&' N2 04 (2)
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UDMH '&'N2 04
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LH2 '&' LOX (3)
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Burn Time (Seconds)
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100
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160
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150
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720
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Maximum Vac Thrust (Kn)
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4736
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680
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720
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73.5
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Specific Impulse (Ns/Kg)
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2610
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2750
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2890
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4510
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Control System
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Multi Port SITVC (4)
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EGC (5) Single plane gimbaling
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EGC two plane gimbaling '&' Hot Gas Reaction Control System
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Two steering engines '&' Cold Gas Reaction Control System
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Data Source: Indian Space Research Organisation (ISRO)
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Legend:
- Hydroxyl Terminated Poly Butadiene (HTPB)
- 75% Unsymmetrical Di-Methyl Hydrazine (UDMH) / 25% Hydrazine Hydrate (mixture prevents combustion instability) and Nitrogen Tetroxide (N2O4)
- Liquid Oxygen and Liquid Hydrogen (LH2 and LOX)
- Multi-port Secondary Injection Thrust Vector Control (SITVC)
- Engine Gimbal Control (EGC)
THE VIKAS LIQUID PROPELLANT ENGINE
THE SAGA OF THE CRYOGENIC UPPER STAGE PROJECT (CUSP)
The story of the cryogenic engine development in India is a case of missed opportunities and lack of appropriate and timely decision-making competencies by ISRO's which failed to foresee the unavoidable need to develop a more powerful final stage to inject heavier communication satellites. ISRO assumed that importing foreign technology would cut short development time. In 1991, ISRO and Glavkosmos of erstwhile Soviet Union signed an agreement to supply two flight worthy stages initially and included an option to supply seven cryogenic stages and one ground mock-up stage, the agreement also included subsequent transfer of know- how to fabricate the engines in India. However, ISRO failed to anticipate the dynamics of international politics and the inevitable collapse of the Soviet Union that ultimately led to the scuttling of the project. A belligerent America under the Missile Technology Control Regime Act, 1987 (MTCR) imposed sanctions on both entities.Though discriminatory in its outlook on its main stated purpose of limiting diffusion of delivery systems capable of carrying nuclear, chemical or biological warheads, MTCR was distinctly a ratification for the Nuclear Club members to continue using and proliferate such technology and prevent technology from going to non club members. However, as per several experts in this realm the Russian cryogenic deal distinctly contravened MTCR provisions. That's an unlikely story which India refuses to accept even now.
However, things did not pan out well for the GSLV program as planned by ISRO, since 2001 there have been multiple unsuccessful attempts. ISRO’s plan was to use the six Russian supplied cryogenic stages for developmental and operational missions and in the interim work on the development of an indigenous cryogenic engine. In April 2010, the GSLV-D3 mission with the first indigenous cryogenic upper stage failed due to a malfunction in the Fuel Booster Turbo Pump. After careful evaluation, ISRO made several modifications in the engine design and the FBTP component to account for the expansion and contraction of bearings and casings, malfunction of a seal location, seizure of rotor and the rupture of turbine and consequently performing various ground tests to prove its technological competence. ISRO also encountered several problems before resolving issues such as production of special alloys, high-speed turbines, fabrication of composite thermal insulation materials, and handling of fluids at very low temperatures.
Other key Design Improvements includes:
- Redesign of Lower Shroud and wire tunnel, which protects the cryogenic engine during atmospheric flight of GSLV-D5
- Modification of Ignition Sequence to ensure the smooth, successful and sustained ignition for Main Engine (ME), Steering Engine (SE) and Gas Generator (GG)
- In addition, indigenisation of many critical systems including Liquid Hydrogen Propellant Acquisition System (to prevent the possibility of contamination), Polyimide pipelines and Liquid Oxygen and Liquid Hydrogen Level Sensors successfully accomplished
WHAT IS A CRYOGENIC ENGINE?
Russian RD-171 - is the world's most powerful Semi-Cryogenic Engine
The Vulcain-2 Cryogenic Engine-ESA - Images: Wikipedia
A cryogenic rocket engine is one that burns a combination of super cooled liquid fuels stored at very low temperatures. Various cryogenic fuel-oxidizer combinations have undergone repetitive tests in the past, but the combination of liquid hydrogen as fuel and liquid oxygen as an oxidizer was found to be the most suitable and efficient. Oxygen liquidizes at –183 degree C and Hydrogen at –253 degree C. Both propellants are easily available and cheap to produce, and when burned it has one of the highest entropy releases by combustion, producing a high degree of specific impulse.
The major components of a cryogenic engine are the combustion chamber or the thrust chamber, pyrotechnic/electric ignitors, fuel injector, fuel turbo pumps, oxidizer pumps, gas turbine, cryogenic valves, regulators, fuel tanks, and the rocket engine nozzle. In terms of feeding propellants to the combustion chamber generally all liquid-propellant engines work as a staged combustion cycle, an expander cycle, a gas-generator cycle or a simple pressure fed cycle.
A cryogenic stage is more efficient than an earth-storable liquid engine for the reason that it provides more thrust for every kilogram of propellant it burns, giving it a substantial payload advantage. However, cryogenic stage is technically very complex to build and operate when compared to a solid or liquid stages due to its use of propellants at extremely low temperatures and the associated thermal and structural problems. It also requires complex ground support systems like propellant storage facilities filling stations, transportation and handling of cryogenic fluids and related safety processes.
INDIAN CRYOGENIC STAGE
ISRO’s Cryogenic Upper Stage (CUS) Project envisaged the design and development of the indigenous Cryogenic Upper Stage to replace the stage procured from Russia and used in earlier GSLV flights. The CE-7.5 main engine and two smaller steering engines of CUS together develop a nominal thrust of 73.55Kn (Kilo Newton) in vacuum. During the flight, CUS fires for a nominal duration of 720 seconds. Liquid Oxygen (LOX) and Liquid Hydrogen (LH2) from the respective tanks are fed by individual booster pumps to the main turbo pump to ensure a high flow rate of propellants into the combustion chamber. Thrust control and mixture ratio control are achieved by two independent regulators. Two gimballed steering engines provide for control of the stage during its thrusting phase. (Text Link: ISRO)
CE-7.5 Cryogenic Engine Stage Shroud - Image: ISRO
ISRO’s Cryogenic Upper Stage (CUS) Project envisaged the design and development of the indigenous Cryogenic Upper Stage to replace the stage procured from Russia and used in earlier GSLV flights. The CE-7.5 main engine and two smaller steering engines of CUS together develop a nominal thrust of 73.55Kn (Kilo Newton) in vacuum. During the flight, CUS fires for a nominal duration of 720 seconds. Liquid Oxygen (LOX) and Liquid Hydrogen (LH2) from the respective tanks are fed by individual booster pumps to the main turbo pump to ensure a high flow rate of propellants into the combustion chamber. Thrust control and mixture ratio control are achieved by two independent regulators. Two gimballed steering engines provide for control of the stage during its thrusting phase. (Text Link: ISRO)
Free to Use
DESIGN IMPROVEMENTS IN THE CRYOGENIC UPPER STAGE
- Modified design of the Fuel Booster Turbo Pump (FBTP), taking care of the expansion and contraction of the bearings and casing at cryogenic temperatures
- Modification of Ignition Sequence to ensure the smooth, successful and sustained ignition for Main Engine (ME), Steering Engine (SE) and Gas Generator (GG)
In addition, indigenisation of many critical systems including Liquid Hydrogen Propellant Acquisition System (to prevent the possibility of contamination), Polyimide pipelines and Liquid Oxygen and Liquid Hydrogen level sensors was successfully accomplished.
In order to validate the design improvements, the following extensive qualification tests were completed at the Main Engine Test (MET) facility and the High Altitude Test (HAT) facility:
- Two acceptance tests for flight unit of FBTP
- High altitude tests to confirm the ignition sequence in flight under vacuum conditions
- Cryogenic Main Engine (200 seconds) and Steering Engine (100 seconds) acceptance tests.All the improvements have been thoroughly reviewed by expert committees including eminent national experts (Text Link: ISRO)
CE-7.5 is a regeneratively cooled, variable thrust, staged combustion cycle engine. In a staged combustion cycle engine some of the propellant is burned in a pre-burner and the resulting hot gas is used to power the engine's turbines and pumps. The exhausted gas is then injected into the main combustion chamber, along with the rest of the propellant, and combustion is completed. Stage combustion engines are more complex in its technology but at the same time more energy efficient.
ISRO formally started the Indigenous Cryogenic Upper Stage Project in 1994. The engine successfully completed the Flight Acceptance Hot Test in 2008, and integrated with propellant tanks, third stage structures and associated feed lines for the first launch. On 27 March 2013, the engine was tested successfully under vacuum conditions at the High Altitude Test Facility, Liquid Propulsion Systems Center, Mahendragiri. The engine performed as expected and was qualified to power the third stage of the GSLV MK-2 rocket. (Text Link: Wikipedia - Modified)
On 5 January 2014, the cryogenic engine performed impeccably and launched the GSAT-14 satellite using GSLV-D5 launcher into a precise GTO orbit. Cryogenic technology has only successfully been developed by a select club of nations which includes the United States, Russia, France, Japan and China. While the technology complexity seems to have been mastered, the real challenge for ISRO will be to sustain the quality and consistency to launch successful GSLV missions like the workhorse PSLV.
On 5 January 2014, the cryogenic engine performed impeccably and launched the GSAT-14 satellite using GSLV-D5 launcher into a precise GTO orbit. Cryogenic technology has only successfully been developed by a select club of nations which includes the United States, Russia, France, Japan and China. While the technology complexity seems to have been mastered, the real challenge for ISRO will be to sustain the quality and consistency to launch successful GSLV missions like the workhorse PSLV.
CE-7.5 Engine Specifications:
- Operating Cycle - Staged combustion
- Propellant Combination - LOX / LH2
- Thrust Nominal (Vacuum) - 75Kn
- Operating Thrust Range - 73.55Kn to 93.1Kn (To be set at any fix values)
- Chamber Pressure (Nominal) - 58 bar
- Engine Mixture ratio (Oxidizer/Fuel by weight) - 5.05
- Engine Specific Impulse - 454 ± 3 seconds (4.452 ± 0.029 km/s)
- Engine Burn Duration (Nominal) - 720 seconds
- Propellant Mass - 12800 kg
NEW UPRATED CRYOGENIC STAGE
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| CE-20 Cryogenic Stage-Image Wikipedia |
The new CE-20 is a cryogenic rocket engine being developed by ISRO to power the upper stage of the next generation heavy-lift booster the GSLV-MK III. The CE-20 is the first Indian cryogenic stage to feature a gas-generator cycle. The engine produces a nominal thrust of 200Kn, but has an operating thrust range between 180Kn to 220Kn and can be set to any fixed values between them. The combustion chamber burns liquid hydrogen and liquid oxygen. The engine has a thrust-to-weight ratio of 34.7 and a specific impulse of 444 seconds (4.35 km/s) in vacuum. (Text Link: Wikipedia - Modified)
FUTURE GSLV – THE GSLV MK III
GSLV MK III referred to as a GSLV-MK2 with "steroids" or the “muscular sibling” is a three Stage heavy lift booster system capable of launching a 10-ton payload in Low Earth Orbit and a 5-ton spacecraft in Geo Synchronous Orbit (GSO). The vehicle will have a lift-off weight of about 630 tons and will be 42.4 meters tall. The First stage will consist of two S200 Large Solid Boosters (LSB) with 200 tons of solid propellant strapped to the second stage and each has a diameter of 3.2 metres and a length of 25 metres. The first stage will have a specific thrust of 5151Kn and will burn for 130 seconds. The second stage will be a restartable 4 meters diameter wide liquid core stage with 110 tons of Unsymmetrical Dimethyl Hydrazine and Dinitrogen Tetroxide liquid propellants, its rated thrust is 1,600Kn with a specific impulse of 300 seconds and a burn time of 200 seconds. This stage will be India's first liquid engine cluster design with two Vikas engines each huddled together to develop a thrust of about 75 tons. The improved Vikas engine will use regenerative cooling, providing improved weight and specific impulse, compared to earlier versions. This stage will be 3.4 meters wide and 25 meters in length with an estimated enhanced thrust of 785 tonne.The rocket will employ an upgraded bigger and more powerful C-25 cryogenic restartable engine. It will have a thrust of 200Kn with a specific impulse of 450 seconds and burn time is estimated at 580 seconds and will use liquid oxygen and hydrogen as fuel. The payload fairing will be voluminous with a 5 meters diameter and payload volumes of 100 cubic meters.Comparable rockets are Angara A3, Ariane 5, Atlas V, Delta IV, Falcon 9 V1.1, Falcon Heavy, H-IIA, H-IIB, Long March 3B, Titan IIIC, Soyuz, Zenit and Proton.
Launch Sequence
The Launch sequence starts with the simultaneous ignition of the two S-200 solid rocket boosters which burn for 130 seconds and at 110 seconds, with the S-200 motors still burning, the core with two clustered Vikas-XL engines (designated L-110) ignites and burns for 200 seconds. Both the S-200 motors are jettisoned at 149 seconds, and at 253 seconds when the vehicle reaches 115 Km the payload fairing is ejected. The L-110 burns out at 311 seconds and separates and the Cryogenic stage C-25 is ignited. The Cryogenic engine burns for 580 seconds with two starts and places the payload in the GTO orbit of 180 X 36,000 Km.
Vehicle Description
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| The Imposing GSLV MK III - Image: ISRO |
GSLV MK III referred to as a GSLV-MK2 with "steroids" or the “muscular sibling” is a three Stage heavy lift booster system capable of launching a 10-ton payload in Low Earth Orbit and a 5-ton spacecraft in Geo Synchronous Orbit (GSO). The vehicle will have a lift-off weight of about 630 tons and will be 42.4 meters tall. The First stage will consist of two S200 Large Solid Boosters (LSB) with 200 tons of solid propellant strapped to the second stage and each has a diameter of 3.2 metres and a length of 25 metres. The first stage will have a specific thrust of 5151Kn and will burn for 130 seconds. The second stage will be a restartable 4 meters diameter wide liquid core stage with 110 tons of Unsymmetrical Dimethyl Hydrazine and Dinitrogen Tetroxide liquid propellants, its rated thrust is 1,600Kn with a specific impulse of 300 seconds and a burn time of 200 seconds. This stage will be India's first liquid engine cluster design with two Vikas engines each huddled together to develop a thrust of about 75 tons. The improved Vikas engine will use regenerative cooling, providing improved weight and specific impulse, compared to earlier versions. This stage will be 3.4 meters wide and 25 meters in length with an estimated enhanced thrust of 785 tonne.The rocket will employ an upgraded bigger and more powerful C-25 cryogenic restartable engine. It will have a thrust of 200Kn with a specific impulse of 450 seconds and burn time is estimated at 580 seconds and will use liquid oxygen and hydrogen as fuel. The payload fairing will be voluminous with a 5 meters diameter and payload volumes of 100 cubic meters.Comparable rockets are Angara A3, Ariane 5, Atlas V, Delta IV, Falcon 9 V1.1, Falcon Heavy, H-IIA, H-IIB, Long March 3B, Titan IIIC, Soyuz, Zenit and Proton.
Launch Sequence
The Launch sequence starts with the simultaneous ignition of the two S-200 solid rocket boosters which burn for 130 seconds and at 110 seconds, with the S-200 motors still burning, the core with two clustered Vikas-XL engines (designated L-110) ignites and burns for 200 seconds. Both the S-200 motors are jettisoned at 149 seconds, and at 253 seconds when the vehicle reaches 115 Km the payload fairing is ejected. The L-110 burns out at 311 seconds and separates and the Cryogenic stage C-25 is ignited. The Cryogenic engine burns for 580 seconds with two starts and places the payload in the GTO orbit of 180 X 36,000 Km.
Vehicle Description
Description
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Specification
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Gross Lift-off Weight
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639 Tonnes
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Gross Propellant Weight
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549 Tonnes
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Total Length
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42.5 Meters
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Payload to GTO
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4 Tonnes
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Total Lift-Off Thrust
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7000Kn
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Total Impulse
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1542 Sec
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Scheduled Flights
Flight
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Launch Date
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Variant
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Payload
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Remarks
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LVM3-X
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May-June 2014
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Mk III
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Crew Module
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Sub-orbital Development flight, passive Cryogenic stage
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E1
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Late 2015-Early 2016
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Mk III
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GSAT-19E
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Orbital first operational flight
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FUTURE BOOSTER - UNIFIED LAUNCH VEHICLE (ULV)
A new series of Satellite Launch Vehicle is on the cards, ISRO is likely to rationalize propulsion stages for the PSLV, GSLV and GSLV MK III in the near future, and this project will focus on a modular approach by incorporating both a cryogenic (C25 - to power the upper stage) and a semi cryogenic stage (SC160). The vehicle will mate solid propellant Strap-on boosters of different variations (S12, S60, S138, and S200) to achieve specific payload requirements. Adopting a common core stage design will facilitate reduction in costs and integration time. The new design philosophy will be essentially based to a great extent on the GSLV MK III configuration.
Liquid propellant used for PSLV and GSLV are toxic and harmful for the environment. The trend worldwide is to change over to eco-friendly propellants. Liquid engines working with cryogenic propellants comprising of liquid oxygen and liquid hydrogen and semi cryogenic engines using liquid oxygen and kerosene are considered relatively environment friendly, non-toxic and non-corrosive, safe to store and handle.
ISRO'S SEMI CRYOGENIC ENGINE
Semi-Cryogenic project envisages the design and development of a 2000Kn semi-cryogenic engine for the future heavy lift launch vehicles and re-usable launch vehicles (RLV). Semi-Cryogenic engine uses a combination of liquid oxygen (LOX) and ISROSENE (Rocket grade kerosene) as propellants, which are Eco-friendly and cost effective. The realization of the engine involves the development of many new critical technologies namely Special Materials and Coatings, Brazing Process, Prevention of Coking, Kerosene Refinement, Seals And Bearings for Turbo Pump, Combustion Instability and Control Components.(Text and Table Link: ISROHQ)
SEMI CRYOGENIC STAGE SPECIFICATIONS
Thrust (vacuum),Kn
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2000
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Isp (vacuum), N-s/kg
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3285
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Chamber Pressure, MPa
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18
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Mixture Ratio
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2.65
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Thrust Throttling (% of nominal thrust)
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60 to 105
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Engine Gimbal
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± 8º in two plane
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CONCLUSION
In conclusion,“ISRO may have stumbled over the cryogenic issue” writes Gopal N Raj, in his amusing book "Reaching for the Stars", “but it remains an organization with considerable vitality. It is one of the few institutions in the country with an excellent record of accomplishment of delivering useful products and services." He further adds.“The cryogenic experience suggests that ISRO needs to change its approach towards the development and acquisition of technology. It will require the ability to initiate technology development sufficiently in advance.” Ironically written 14 years ago in 2000 his views stand in good stead even today.
Further Reading:
- Reaching for the Stars - Gopal N Raj - First published in Viking by Penguin India 2000
- Various pages, publications and newsletters of ISRO and ISROHQ - which can be found in their respective portals - http://www.isro.org/ and http://isrohq.vssc.gov.in/
