GSLV Mk. II Launch Vehicle
The Geosynchronous Satellite Launch Vehicle, better known by its abbreviation GSLV, is an Indian expendable launch system that was developed and is operated by the Indian Space Research Organization.
The GSLV project was initiated back in the 1990s when India determined that it needed its own launch capability for Geosynchronous Satellites to become independent from other launch providers. At the time, India was relying on Russian/Soviet launch vehicles for heavy satellite launches. With the emergence of commercial launch providers, such as Arianespace, India shifted its GSO Satellites to those while GSLV was being developed.
The launch system uses a large number of heritage components already employed on the Polar Satellite Launch Vehicle that first flew in 1993. The three-stage GSLV has an improved performance over four-stage PSLV with the addition of strap-on liquid-fueled boosters and a cryogenic upper stage. GSLV uses a combination of solid fueled, liquid-fueled and cryogenic stages. The vehicle weighs 414,000 Kilograms at liftoff standing 49 meters tall with a core diameter of 2.8 meters.
The first stage is the S139 solid-fueled stage that is also used on PSLV. Around the core, four strap-on liquid-fueled boosters are mounted each featuring a Vikas engine using storable propellants. The second stage is also a storable propellant stage using a single modified Vikas engine while the third stage is a cryogenic stage using liquid Oxygen and liquid Hydrogen that is consumed by an ICE engine. The vehicle can deploy payloads of up to 2,500 Kilograms to a Geosynchronous Transfer Orbit, Low Earth Orbit Capability is 5,000kg. GSLV is operated from the Satish Dhawan Space Center.
Over the course of its development, GSLV flew in various configurations being designated Mk I a, b and c, and Mk II with Mk III being the successor to the first generation of GSLV launchers scheduled to make its first flight in 2014.
In the Mk Ia configuration, the most basic version of the launcher that is now retired, GSLV used a 125-ton Core Stage and a Russian-built Cryogenic Upper Stage since the Indian-developed cryogenic stage required more time to be designed and built. Mk Ia flew its first development flight on April 18, 2001 marking the first launch of the GSLV class vehicle. The flight was only a partial success as the launcher delivered its payload, the GSAT-1 Communications Satellite, to a lower-than planned orbit due to a shortfall in performance either caused by the vehicle’s guidance system or a premature shutdown of the third stage. GSAT-1 ended up in a lower orbit and due to a design flaw in its propulsion system, was unable to reach Geostationary Orbit rendering its useless for its original purpose.
In May 2003, Mk Ia flew for the second time. On that mission, GSLV performed as planned and successfully delivered the GSAT-2 payload to Geosynchronous Transfer Orbit. The next GSLV launch came in September 2004 and was the first flight of the Mk Ib configuration that still used the Russian upper stage, but featured the 139-ton first stage. The flight was a success and delivered the GSAT-3 spacecraft to its intended orbit to serve as experimental communications satellite.
On its next flight in July 2006, GSLV suffered another failure. Shortly after launch, the vehicle had to be destroyed by Range Safety Personnel because it veered off its pre-planned course due to the failure of one of the boosters. Remains of the rocket and the INSAT-4C payload fell into the Bay of Bengal.
GSLV Mk Ib flew again in September 2007 – successfully reaching orbit, but placing the INSAT-4CR satellite into a lower-than-planned orbit at a higher inclination due to a guidance system issue. The spacecraft reached its orbit using its own propulsion system and became fully operational, making this mission a partial success.
The sixth flight of the GSLV marked the first flight of the Mk II variant that uses the Indian-built cryogenic upper stage. The flight test was not successful as the rocket encountered a malfunction of the Fuel Booster Turbo Pump on its third stage causing the loss of the vehicle and GSAT-4 satellite. Flight 2 of the Mk II version was attempted in 2010 and marked another failure as the vehicle was destroyed by the Range Safety Officer after a loss of control that was the result of a structural failure.
After that, the GSLV launch system underwent a thorough review and improvements were made to its Guidance System and Upper Stage to increase its reliability. The GSLV Return to Flight Mission was successfully performed in 2014.
| Type | GSLV |
| Height | 49m |
| Diameter | 2.8m |
| Span | 6.9m |
| Launch Mass | 414,000kg |
| Stage 1 | S139 |
| Boosters | 4 x L40 |
| Stage 2 | GS2 |
| Stage 3 | GS3 |
| Mass to LEO | 5,000kg |
| Mass to GTO | 2,500kg |
GSLV Specifications
The Geosynchronous Satellite Launch Vehicle in its Mk II configuration stands 49 meters tall, with a Core Diameter of 2.8 meters and a liftoff mass of about 414,000 Kilograms.
The vehicle features three stages plus an optional fourth stage. The first stage is a solid-fueled stage holding 138,000kg of propellant.
Around the Core Stage, four strap-on, liquid-fueled boosters are installed. One of the oddities about GSLV is that the four boosters burn longer than the Core Stage does. The second stage of the vehicle is liquid-fueled and uses storable propellants.
The upper stage is a cryogenic stage that uses LOX and LH2 as propellants.
First Stage
| Type | S139 |
| Inert Mass | 28,300kg |
| Launch Mass | 166,300kg |
| Diameter | 2.8m |
| Length | 20.13m |
| Case Material | Maraging Steel |
| Propellant | Solid – HTPB Based |
| Propellant Mass | 138,000kg |
| Guidance | From Upper Stage |
| Propulsion | S139 Solid Rocket Motor |
| Thrust (Vacuum) | 4,860kN |
| Impulse | 105sec |
| Burn Time | 106.9sec |
| Attitude Control | via Boosters, SITVC (Optional) |
| Stage Separation | Flexible Linear Shaped Charge |
| Hot Staging |
The Core Stage of the GSLV is called S139 and is derived from the PS1 Core Stage of the Polar Satellite Launch Vehicle. The only change is the removal of the Secondary Injection Thrust Vector Control System that is needed on PSLV. On GSLV, the Boosters are used to control the vehicle during first stage flight, thus eliminating the need for a Thrust Vector Control System on the core stage. Adding SITVC onto the GSLV is optional and was only performed on its very first flight.
The S139 has an inert mass of 28,300kg and holds 138,000kg of HTPB-based (Hydroxyl-terminated polybutadiene) solid propellant. The stage is 20.1 meters long and 2.8 meters in diameter featuring a maraging steel case. It has a vacuum thrust of 4,860 Kilonewtons (495,600kg) and burns for 107 seconds.
The stage separates with the four boosters once they are reaching depletion. Staging between S139 and GS2 (Stage 2) is accomplished in hot-staging mode – the second stage ignites 1.6 seconds ahead of Booster Shutdown. When the boosters have shut down, the two stages are separated by flexible linear shaped charges that pyrotechnically separate the two stages allowing the spent first stage and boosters to be pushed away by the second stage. This maneuver comes at the cost of propellant and performance but minimizes propellant unsettling that occurs when igniting in coast mode.
First Stage Nozzle-End Segment
Boosters
| # Boosters | 4 |
| Type | LH40 |
| Length | 19.7m |
| Diameter | 2.1m |
| Inert Mass | ~5,600kg |
| Launch Mass | 47,600kg |
| Tank Material | Aluminum Alloy |
| Fuel | UH25 – 75% UMDH, 25% Diazane |
| Oxidizer | Nitrogen Tetroxide |
| Propulsion | 1 Vikas 2 |
| Thrust | 763kN |
| Impulse | 293 sec |
| Engine Dry Weight | 900kg |
| Engine Length | 2.87m |
| Engine Diameter | 0.99m |
| Burn Time | 148sec |
| Chamber Pressure | 58.5bar |
| Mixture Ratio | 1.7 (Ox/Fuel) |
| Attitude Control | Single-Plane Engine Gimbaling |
| Stage Separation | With Core Stage |
Four liquid-fueled boosters are clustered around the Core Stage of the vehicle. Each is 2.1 meters in diameter and 19.7 meters long facilitating two Aluminum propellant tanks that can hold about 42,000 Kilograms of Nitrogen Tetroxide Oxidizer and UH25 fuel – a mixture of 75% Unsymmetrical Dimethylhydrazine and 25% Hydrazine Hydrate. (UH 25 prevents combustion instability)
Each Booster is equipped with a single Vikas 2 engine which is a Viking engine that was used aboard the European Ariane 1 launcher and is now manufactured under license in India. The Vikas engine used on GSLV is a lightly modified Viking 2 engine. It is 2.87 meters long and 0.99m in diameter and weighs 900 Kilograms. The engine operates at a Chamber Pressure of 58.5 bar and uses an Oxidizer to Fuel Ratio of 1.7. Vikas 2 delivers a thrust of 763 Kilonewtons (77,800kg). The four boosters have a burn time of 148 seconds.
The Vikas engines on each booster can be gimbaled in a single plane allowing three-axis control during first stage flight.
The four boosters ignite 4.6 seconds prior to the first stage to allow the Vikas engines to reach operational conditions before the Core Stage is ignited and the rocket blasts off. In flight, the four boosters continue to burn after first stage shutdown and are separated from the vehicle with the first stage. The advantage of this simpler design is that a Booster Separation event is avoided, but it comes at the cost of performance because the four boosters have to propel the first stage once it has burned out which represents nearly 30 tonnes of dead weight.
| Type | GS2 – L37.5H |
| Inert Mass | ~5,500kg |
| Launch Mass | 44,900kg |
| Length | 11.56m |
| Diameter | 2.8m |
| Tank Material | Aluminum Alloy |
| Fuel | UH25 – 75% UMDH, 25% Diazane |
| Oxidizer | Nitrogen Tetroxide |
| Propellant Mass | 39,400kg |
| Guidance | From Upper Stage |
| Propulsion | 1 Vikas 4 |
| Thrust (Vacuum) | 799kN |
| Impulse | 293s |
| Engine Dry Weight | 900kg |
| Engine Length | 3.51m |
| Engine Diameter | 1.70m |
| Burn Time | 158sec |
| Chamber Pressure | 58.5bar |
| Mixture Ratio | 1.7 (Ox/Fuel) – MR Optimization |
| Area Ratio | 31 |
| Prop Flow Rate | 278.04kg/s |
| Attitude Control | Main Engine Gimbaling, Roll RCS |
| Stage Separation | Merman Band |
Second Stage
The second stage of the GSLV launcher, designated GS2, also uses hypergolic propellants – NTO and UH25. It has a launch mass of 44,900kg being 11.6m long and 2.8m in diameter.
The tanks are made of Aluminum alloy and hold 39,400kg of storable propellants that are being consumed by a single Vikas 4 engine.
The engine is based on the Viking 4 of the Ariane 1 launcher and also features slight modifications. It is optimized for operation in vacuum conditions with an extended nozzle that has an area ratio of 31.
The engine is 3.51m long and 1.7m in diameter weighing about 900kg.
It provides 799 Kilonewtons of vacuum thrust (81,500 Kilograms) over the course of its 158-second burn. It operates at an Ox. to Fuel Ratio of 1.7 that can be optimized by the Flight Control System.
Vehicle Control during the second stage burn its provided by gimbaling the main engine by up to 4 degrees for pitch and yaw. Roll Control is provided by a Cold Gas Thruster System.
The second and third stage again separate in a hot-staging mode – the second stage shuts down and at the same time, the ignition of the third stage and the stage separation mechanism, a Merman Band Sep System, are initiated.
Third Stage
| Type | GS3 – C15 |
| Inert Mass | ~2,500kg |
| Launch Mass | ~15,300kg |
| Length | 8.7m |
| Diameter | 2.8m |
| Tank Material | Aluminum Alloy |
| Fuel | Liquid Hydrogen |
| Oxidizer | Liquid Oxygen |
| Propellant Mass | 12,800kg |
| Guidance | Inertial Platform, Closed-Loop |
| Propulsion | 1 ICE (CE-7.5) |
| Cycle | Staged Combustion |
| Thrust (Vacuum) | 73.5 to 93.1kN |
| Impulse | 454sec |
| Engine Dry Weight | 435kg |
| Engine Length | 2.14m |
| Engine Diameter | 1.56m |
| Burn Time | Up to 1,000sec |
| Chamber Pressure | 58bar |
| Attitude Control | 2 vernier Jets, each 2kN |
| RCS for Coast Phases | |
| Stage Separation | Merman Band, Hot Staging |
The third stage or GS3 of the GSLV Mk II is an Indian-built cryogenic upper stage. It is 8.7 meters long and 2.8 meters in diameter featuring two Aluminum Alloy Tanks that hold about 12,800 Kilograms of Liquid Hydrogen and Liquid Oxygen. The inert mass of the third stage is about 2,500kg.
It is powered by a single ICE (Indian Cryogenic Engine) or CE-7.5. The engine is a staged combustion type engine.
Some of the propellant is used to power the turbopump of the engine before being injected into the main combustion chamber along with the rest of the propellant. The turbopump spins at about 42,000 rpm.
The engine weighs 445 Kilograms and is 2.14 meters long and 1.56 meters in diameter operating at a chamber pressure of 58bar. It provides a nominal thrust of 73.5 Kilonewtons (7,500kg), but can be throttled up to 93.1kN (9,500kg).
Usually, the engine operates at a higher thrust level for the first 300 seconds of its burn before throttling down to nominal thrust for the remainder of its firing that can be up to 1,000 seconds in duration.
Vehicle control is provided by two vernier jets that can be swiveled in all directions to provide three-axis control.
Each vernier provides 2kN (204kg) of thrust. During Coast Phases, a cold gas reaction control system is used for vehicle stabilization and re-orientations.
The third stage can be re-ignited in flight. The spacecraft is separated by spring thrusters mounted at the separation interface with the third stage.
The third stage of the launch vehicle also houses the flight computers and the inertial guidance platform of the GSLV.
The control system was developed and built in India. GSLV uses a Redundant Strap Down Inertial Navigation System/Inertial Guidance System that is housed in the equipment bay of the third stage. The digital control system of the launcher uses closed-loop guidance throughout the flight to ensure accurate injections into the target orbit.
Also mounted on the third stage is the communications system of the launch vehicle consisting of an S-Band system for telemetry downlink and a C-Band transponder that allows radar tracking and preliminary orbit determination.
The communications link is also used for range safety / flight termination that uses a dedicated system that is located on all stages of the vehicle and features separate avionics.
Payload Fairing
| Diameter | 3.4m |
| Length | 7.8m |
| Construction | Aluminium Alloy |
| Sep Altitude | 115km |
The Payload Fairing, or “Heat Shield” as ISRO refers to it, is positioned on top of the stacked vehicle and its integrated Payload. It protects the spacecraft against aerodynamic, thermal and acoustic environments that the vehicle experiences during atmospheric flight. When the launcher has left the atmosphere, the fairing is jettisoned by pyrotechnical initiated systems. Separating the fairing as early as possible increases launcher performance.
The fairing of the GSLV is 3.4 meters in diameter and 7.8 meters in length offering enough space for a variety of payloads that are in the weight-category of GSLV. The fairing is made of Aluminum Alloy featuring acoustic absorption blankets.
The fairing is separated at an altitude of 115 Kilometers. Separation is accomplished by a linear piston cylinder separation and jettisoning mechanism (zip cord) running along the full length of the PLF and a clamp and joint at the base of the fairing. Both systems are pyrotechnically initiated. The gas pressure generated by the zip cord expands a rubber bellow that pushes that piston and cylinder apart, pushing the fairing halves laterally away from the launcher.
Optional Fourth Stage
The GSLV launcher can be outfitted with a fourth stage that could serve as apogee module to boost a satellite or spacecraft into its final orbit or trajectory. Usually, satellites and spacecraft are equipped with their own propulsion systems that are capable of performing apogee maneuvers so that the fourth stage is not required.
A possible fourth stage design for GSLV closely resembles the fourth stage of the PSLV launcher that uses storable propellants to provide precise injection capability.
