Indian Cryogenic Engine: The Story of Four Decades

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While the indigenisation of a technology acquired from Russia has taken two decades, ISRO’s efforts at developing an entirely indigenous technology are four decades old. By R. RAMACHANDRAN
The successful launch of the developmental flight D5 of the Geosynchronous Satellite Launch Vehicle (GSLV) on January 5 marks the end of a 20-year-long and difficult journey to the indigenisation of the Russian cryogenic technology acquired in 1995. The upper stage of this three-stage GSLV was powered by the indigenously built cryogenic engine and stage, which was modelled entirely on the Russian engine KV1 designed to deliver 7.5-tonne thrust (75 kN) and the Russian stage 12KRB designed to carry 12.5 tonnes of liquid propellants, liquid oxygen (LOX) and liquid hydrogen (LH2).

When Glavkosmos, the commercial arm of the Russian space agency, reneged on the agreement on transfer of cryogenic engine technology to the Indian Space Research Organisation (ISRO) in 1993, U.R. Rao, the then ISRO Chairman, famously claimed to the media at various fora that since some data, drawings and information had already been received from the Russians, indigenous development based on that technology would be achieved by 1997. In off-the-record conversations, however, ISRO scientists admitted that it would take a minimum of 10 years. Retrospectively, even that turned out to be highly optimistic. It has taken a full two decades.

While indigenisation of an acquired technology has taken two decades, ISRO’s efforts at an entirely indigenous development of the technology are four decades old. Once import of Russian technology became possible, indigenous development, which, since the 1970s, has been progressing in fits and starts (see separate article on page 22) right up to the new millennium, seems to have given way to an indigenous development route that has been strongly influenced by the acquired technology. (The indigenised GSLV is called GSLV-MkII as against GSLV-MkI, which uses the Russian engine and stage.) It is moot whether this was the right approach to develop a complex technology given the extraordinarily long time that the indigenisation process has taken. The truly indigenous GSLV-MkIII, powered by an engine and stage that are totally indigenous, awaits its test next year.

As recounted by S. Ramakrishnan, former Director of ISRO’s Liquid Propulsion Systems Centre (LPSC) and the current Director of the Vikram Sarabhai Space Centre (VSSC), in his Aeronautical Society of India presentation in June 2012, as part of the fully indigenous developmental efforts that began in the 1980s, many ground tests were carried out on a one-tonne thrust subscale engine in order to generate data for designing a large-scale 12-tonne thrust engine (C12).

The first experimental hot test (real life test burning with the actual propellant) on the subscale engine was conducted in 1987 using gaseous hydrogen (GH2) and gaseous oxygen (GO2). This was followed by a test with the propellant combination of LOX and GH2 that used a heat-sink thrust chamber. The first fully cryogenic regeneratively cooled subscale engine, which used LOX and LH2, was tested after a five-year gap, in June 1993, at a newly built subscale test facility. In a regeneratively cooled engine, one of the propellants from the engine outlet passes around the chamber to prevent it from overheating.

The test had to wait for the setting up of the first ever liquid hydrogen plant in the country, which ISRO did. Also the revival of the indigenous development that was hitherto on a slow mode was spurred following the uncertainty over the Russian deal. The test was a failure, which was caused by the failure of an isolation valve (also called the “latch”), an important component of the test-firing equipment, resulting in a hydrogen explosion. The test had apparently been attempted using a valve unsuitable for cryogenic liquid firings. Identification of the problem led to the use of an Oxygen Free High Conductivity (OFHC) copper, instead of stainless steel (SS), thrust chamber. A subsequent test in November 1995 after modifying the test procedure was, however, successful, which was followed by more tests.

According to Ramakrishnan, a sea-level thrust chamber for C12 was designed, realised and two hot tests in de-rated conditions, using LOX and LH2, were carried out in 1998. Experimental studies were also conducted on turbo pump systems. These were tested with gaseous nitrogen as turbine drive fluid and pump handling water in May 2000. “These studies,” says Ramakrishnan in his paper, “gave the ISRO team the experience on design, analysis and realisation of cryo thrust chambers and turbo pumps, as well as in production, storage and handling of cryogenic propellants.” The regeneratively cooled subscale engine incorporated some important technology developments, like electroforming where thin metallic layers are deposited on a base surface by electrodeposition. Electroforming has been used in the French HM7, the main space shuttle engine, and the Japanese LE7 to form the coolant channels through which one of the propellants would pass around the engine. The electroforming technique for the one-tonne cryogenic subscale engine (and later for the 12-tonne (C12) engine as well) had been developed by the CSIR’s Central Electrochemical Research Institute (CECRI) at Karaikudi, Tamil Nadu.

The Russian engine, on the other hand, uses the vacuum brazing technique to fabricate the coolant channels and, for indigenising this aspect of the engine, a large, power-guzzling, brazing apparatus has been imported and the facility established at Godrej, Mumbai, because of which the indigenous development of electroforming technology seems to have been given the go by. According to K.N. Srinivasan of CECRI, a 12-tonne electroformed thrust chamber had been supplied to ISRO for testing at the Mahendragiri (southern Tamil Nadu) liquid engine testing complex. However, he was not aware of what the status of the test was and whether the technology will be used by ISRO at all.

When asked about this, Ramakrishnan said in an email response: “We had realised a 12-tonne thrust chamber by electroforming route and tested it at about 60 per cent thrust level after which we stopped the programme. However, for GSLV-Mk III, the 20 t thrust engine (CE20) development is progressing well. Of course, C12 experience has helped. However, the chamber for CE20 is a brazed version based on Russian engine technology. We are not pursuing electroforming route anymore.” Having invested in a brazing apparatus, ISRO has perhaps decided in favour of the brazing route.

The ISRO-Glavkosmos contract (see article on page 22) provided for seven cryogenic stages. The stage had a propellant (LOX+LH2) loading of 12.5 tonnes and was powered by an engine designed to deliver a thrust of 7.5 tonnes. This engine, originally designated as RD-56, had been developed for the Soviet moon mission in the 1960s. This was modified for use in the GSLV, and the stage too was specially designed for the GSLV. Besides the main engine, the propulsion unit had two pressure-fed vernier (steering) engines as well, which could provide attitude control by swivelling or gimballing. The engine used the complex staged combustion cycle (SCC) as against the simpler and more flexible gas generator cycle (GGC).

Significantly, both the subscale and the C12 engine were based on the GGC. “The C12 cryo engine development programme,” says Ramakrishnan, “was abandoned once the Russian contract was signed to get cryo technology and right now no work on C12 is going on. The focus changed to indigenous version of Russian cryo engine which incidentally operated on SCC unlike GGC proposed for C12.” Interestingly, however, for the GSLV-MkIII, powered by the indigenously designed 20-tonne thrust CE20 engine (and C25 stage) with a much greater target payload capacity of four tonnes-plus, ISRO scientists have returned to the GGC-based thrust generation. (Also see interview with ISRO Chairman K. Radhakrishnan.)

According to Ramakrishnan, the contract provided for only the propulsion hardware to be supplied by Glavkosmos. The control systems required for the stage and mission management—including sequencing, tank pressure control, thrust and propellant mixture ratio control, gimbal control and post-flight passivation, the facilities for stage preparation and propellant servicing—which have provided the much-needed experience in the indigenisation process were to be developed by ISRO. Further, during the development and qualification of the engine and stage for carrying out engine hot tests in Russia, ISRO’s avionics were used for thrust and mixture control and steering engine gimbal control. Stage-level cold flow and hot tests were also conducted in Russia using ISRO’s GSLV equipment bay.

Stage test in Russia

According to the authors of the book A Brief History of Rocketry in ISRO, a stage test in Russia was not part of the original contract. It had been hoped that necessary facilities would be ready at Mahendragiri within the projected time frame for these tests. But there was a delay in getting these facilities ready. Therefore, an additional contract was concluded for stage tests in Russia, presumably at a higher price than it would have cost otherwise.

Following the cancellation of the technology transfer part of the deal, ISRO proposed a Cryogenic Upper Stage (CUS) project, which was approved in April 1994. The objective of the project was to realise an indigenous Cryogenic Stage with the same specifications, configuration and interfaces as the Russian CS for the continuance of operational GSLV flights after all the stages supplied by Glavkosmos were flown.

“The 2.8 m diameter Indian CUS, with 73.5 kN [~7.5 t] nominal thrust SCC engine incorporated several changes from the Russian CS at component and sub-system level, such as pressure regulation and pyrogen igniters, based on the expertise and experience at ISRO,” points out Ramakrishnan in his paper. “However, the main stage architecture of fixed main engine with two steering engines was adopted from CS as such.”

For cryogenic operations, new facilities had to be created. ISRO established LH2 and LOX production plants at Mahendragiri. An integrated LH2 plant at 500 kg/d capacity was set up under a contract with Linde AG, Germany, in 1995. Since cryogenic filling operations before the launch are a great deal more complex than filling operations for earth-storable liquid propellants such as UDMH and NO, a separate contract with Glavkosmos was arrived at for establishing the cryogenic propellant filling plant at the SHAR launch complex. For fabrication of engine and stage hardware of the main engine and stage systems, infrastructure was established at Godrej, and MTAR, Hyderabad, and Hindustan Aeronautics Limited, Bangalore.

The first indigenous engine, A0 was tested on February 16, 2000. The test failed after firing for only 13.7 seconds, and when it was aborted by manual intervention as a fire was detected around the engine ion test stand. The engine and hardware were totally damaged. The first test with A1 was terminated prematurely too as the ignition did not take place in the gas generator. Following a failure analysis, the pyrotechnic igniter was replaced by a pyrogen igniter. Following a series of nine tests on A1, accompanied by several problems, changes and modifications of the configuration, the first successful test of A1 happened on February 9, 2002.

A2, whose configuration is similar to A1 but for a few additional subsystems, underwent 10 tests, including an endurance test that lasted for 1,000 s. The first vernier engine was successfully tested at the subscale test facility using GH2 and LOX on June 25, 2001. According to Ramakrishnan, after extensive development and qualification tests at engine level, the indigenous stage was realised and subjected to successful hot test for the flight duration of 720 s on November 15, 2007. The tests confirmed that the indigenised engine was capable of delivering 27 per cent more thrust than the Russian version.

Indigenous flight stage

The first indigenous flight stage flew in the developmental flight GSLV-D3 on April 15, 2010. The engine fired successfully but could not develop the requisite thrust owing to a malfunction of the LH2 turbo pump. Before the first successful launch of D5 took place, nearly 40 hot tests on seven main engines for a cumulative duration of nearly 8,000 s and about 80 firings on 18 steering engines for a cumulative duration of about 10,500 s had been carried out.

The idea to replicate the Russian design for GSLV-MkII was done with a view to minimise the development time and lessen the risks for early operationalisation of the GSLV. From that perspective, even the special materials required for the engine and stage were procured from Russia by 1996 and successfully indigenised. Also, the six launches with the Russian stages, of which four were failures, however, have given ISRO scientists the necessary lessons for successful indigenisation (see interview with K. Radhakrishnan).

These launches—successful, sub-optimal and failed— also provided the necessary experience and data for precise calculation of aerodynamic loads on the upper cryogenic stage, optimisation of the launch azimuth and “range safety” considerations, increasing the thrust rating, higher propellant loading in the solid booster and the liquid stages, increasing the pressure of the liquid stages L40 and L37.5, reduction in the mass of the equipment bay and use of additional composite materials where possible, and obtaining optimal performance of the vehicle in terms of its payload capacity against the targeted delivery of a 2.5-tonne satellite in the Geosynchronous Transfer Orbit (GTO).

During the initial interactions with the Russian engineers, with the best values for the aerodynamic loads and the optimal launch azimuth given the “range safety” constraints (which meant unspent fuel in the second stage liquid engine L37.5), the mass of the cryogenic stage had increased considerably from the original indications and it seemed that the satellite mass could not exceed 1.5 tonnes. Indeed, the satellite GSAT-1 aboard the first developmental launch GSLV-D1 weighed only 1,540 kg.

But later exercises of optimisation resulted in the weight of EDUSAT, launched by the first operational flight GSLV-F01, being increased to 1,950 kg. In the failed launch GSLV-D3, which used the indigenous cryo engine and stage for the first time, the weight of the satellite GSAT-4 could, however, be increased to 2.2 tonnes. In the recent successful launch with the Indian cryo, the weight of GSAT-14 was kept somewhat less at 1,980 kg.

“The first Indian cryo stage which flew in GSLV-D3,” says Ramakrishnan, “was indeed optimised and was lighter than the Russian cryo stage. Also, in D3 we operated the engine at an uprated thrust level. Another factor was more favourable launch azimuth. All these pushed up the payload to 2.2 tonnes. Of course in D5, we played it more conservative. On the basis of shroud failure experience from F06 flight, we made the structure more robust. Also in general, everywhere we kept more margins since our main objective was to prove the working of the Indian cryo stage. We can progressively enhance the payload to 2,250 kg and maybe ultimately 2,500 kg in the coming flights” (see interview with Ramakrishnan).

The next phase of cryogenic technology development, following the expertise gained in the CUS project as well as the successful launch of GSLV-MkII, is the development of GSLV-MkIII, to be powered by a 20-tonne thrust cryo engine CE20 based on GGC. It is expected to have a specific impulse of 443 s. “GGC was deliberately chosen,” writes Ramakrishnan in his paper, “to facilitate faster development by way of individual subsystems testing and qualification before integrated engine test.”

According to him, the various subsystems have been completed and the gas generator, injector, LH2 and LOX turbo pumps, start-up systems and so on have undergone development tests. The integrated power head of the engine developing a turbine power of 2 MW was successfully tested in bootstrap mode on July 30, 2010. “Indigenously designed turbo pump unit of this power level to be successfully realised within the country marks a major milestone in the area of aerospace technology in India,” he says.

The CE20 and C25 are entering the crucial stage for validation tests, which are aimed at the launch of GSLV-MkIII next year. The immediate target for MkIII, however, is a successful launch of the experimental LVM3-X mission in the next few months. LVM3-X will be a GSLV-MkIII rocket except for a passive cryo stage with non-functional cryo engine to test all other elements of this new indigenous ISRO beast.

Despite the recent success of the GSLV and an upbeat ISRO looking into the future with an entirely indigenously developed cryogenic engine, which no doubt will have a mix of the legacy of the Russian engine and domestic efforts of the 1980s and the 1990s, it is still worth asking if the long-drawn indigenous developmental route via Moscow was the best way of achieving it.

Russian route | Frontline

 

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A SUBSTANTIAL PART of the credit for India’s cryogenic success should go to S. Ramakrishnan, Director of the Vikram Sarabhai Space Centre (VSSC), Thiruvananthapuram. He was until December 2012 Director of the Liquid Propulsion Systems Centre (LPSC) where the indigenous cryogenic engine and stage for the GSLV-D5 were built and tested.

After Ramakrishnan joined ISRO in August 1972, he worked as a member of the SLV-3 project team under the leadership of A.P.J. Abdul Kalam. Subsequently, he joined the Polar Satellite Launch Vehicle (PSLV) programme and was Project Director of several PSLV flights from 1996 to 2002. He was also Mission Director of the first four operational flights, PSLV-C1 to PSLV-C4. He played an important role in developing the PSLV’s liquid propulsion stages and in the vehicle’s integration. Under his leadership, the PSLV has become ISRO’s trusted workhorse and the weight of the satellites it can put into orbit went up from 900 kilograms to 1,500 kg.

He graduated in Mechanical Engineering from the College of Engineering, Guindy, Madras (now Chennai), in 1970, and took his M.Tech in Aerospace from Indian Institute of Technology Madras with first rank.

Excerpts from an extended interview Ramakrishnan gave Frontline at Sriharikota and followed up on email:

How do you assess the significance of the GSLV-D5 mission with India’s own cryogenic stage?

It is a very important mission. When we started 20 years ago, we had an agreement with the Russians [to receive the cryogenic technology and stages from them]. But it got thwarted by the Mission Technology Control Regime and we could not get the technology. So it was a difficult job for us to absorb the technology when we started 20 years ago. We had started working on cryogenic rockets in small steps from 1985 onwards. Handling and processing of procured stages [from Russia] gave us the leapfrogging benefit in this advanced technology.

When we started indigenising cryogenic technology, industries had to realise the welded thrust chamber and the high-speed rotating machines. Initially, there were defects and rejections in the hardware coming from the industries. The situation stabilised. The engine testing was done for a total duration of 6,000 seconds and we built the indigenous cryogenic stage and did the full-duration ground test.

Then we launched the GSLV-D3 on April 15, 2010. The starting conditions for the cryogenic engine in that flight were good. Unfortunately, the fuel booster turbo pump [FBTP] stopped immediately after it started. We had to go into all the possible failure modes and study each one of them. At Mahendragiri, we set up the High Altitude Test [HAT] facility and tested the engine ignition and the start process for the first three seconds. So, we were confident this time and we are happy that all our analyses were right. The indigenous cryogenic stage worked well. The injection was precise. The apogee and perigee errors were very small.

You had mentioned earlier that in rocket propulsion technology, cryogenic technology is the most complex.

Yes. Cryogenic technology is the most complex of propulsion technologies. We are the sixth country after the United States, Russia, France, Japan and China to have the cryogenic engine technology. Today, we have mastered it.

We are able to make our own cryogenic stage and fly it. It gives us the confidence to go for the bigger and more powerful GSLV-Mk III, where we have got a totally indigenous 20-tonne-thrust engine with propellants weighing 25 tonnes. The entire vehicle weighs 600 tonnes. It is a 50-metre-tall, three-stage vehicle.

The next GSLV flight with an indigenous cryogenic engine can be within a year. The GSLV has been identified for Chandrayaan-2. Once we make one more successful launch, we can confidently assign it for Chandrayaan-2. We are sure we will be able to repeat the success.

Was there any technology transfer from Russia before the U.S. forced it to renege on the technology transfer agreement in 1993?

There was no technology transfer. We had a lot of ambiguities and uncertainties in the documents. It was a totally new technology when we started. We were confident when we went in for the GSLV-D3. But we faced a failure in it. We took a lot of time to decipher why the FBTP did not sustain. We went into the failure modes systematically. We gave the Failure Analysis Committee’s findings to a national committee. Then we studied all the past GSLV flights from 2001.

In the last three years, we did more tests on the engine subsystems and vacuum ignition at the HAT facility and we did acceptance tests on the individual components such as booster pump, turbo pump, etc. The integration process was streamlined. Cleanliness control was done. We made sure there was no room for another failure.

We knew we had the right technology and the right approach in realising the fabrication, assembly and testing of the indigenous cryogenic engine and that we can make it more robust. Like the PSLV, the GSLV will be a reliable vehicle.

Can you recall ISRO’s early struggle in realising the sub-scale cryogenic engine and later the flight version, especially after the U.S. pressured Russia not to transfer cryogenic technology to India?

Consequent to the cancellation of the cryogenic technology transfer agreement by Russia, ISRO constituted the cryogenic upper stage [CUS] development project in 1995 with the objective of indigenously realising a cryogenic stage with the same specifications, configuration and interfaces as the Russian cryo stage [CS] to sustain the GSLV operational flights after the Russian-supplied cryo stages were used up. In fact, the Indian CUS was aimed as a one-to-one replacement of the Russian CS. Though there were several changes at the stage systems level from the Russian CS, the engine itself was identical to the Russian engine and was getting fabricated at Indian industries based on drawings generated from the data/documents which we could get from the USSR [Union of Soviet Socialist Republics] during the initial phase before the technology transfer agreement was annulled and the information flow stopped.

There was no subscale version and the very first engine assembly realised from Indian industries was the full-scale, 73.5 kN [kilonewton]-thrust staged combustion engine, including the pair of pressure-fed vernier engines. Of course, the total pyro-initiated elements of the engine, that is, the pyro valves and the igniters, were Indian.

The first engine assembly [A0] realised indigenously blew up during the ground test because of an open flame caused by leaking hydrogen cutting the command line leading to the closure of liquid hydrogen [LH2] supply to the engine.

This failure was definitely demoralising and it took considerable time and detailed analysis to reconstruct the sequence of events leading to this catastrophic failure with some collateral damage to the test stand also.

However, we could quickly realise the next engine assembly, A1, and successfully test it, leading to many more long-duration tests at the Mahendragiri facilities of the LPSC.

There was some cryogenic technology transfer from Russia to India in 1992 and 1993 before the tap turned dry. Why did it take 18 years to realise a flight-worthy cryogenic stage, which was used in the April 2010 GSLV-D3 flight?

When we talk about technology transfer for a complex engineering system such as a cryogenic rocket, we have to understand that it is an involved process which takes time and very intense interaction between the technical experts of the two agencies involved. Large volumes of data, documents and drawings have to be exchanged with concurrent interaction to clarify technical points, issues and ambiguities/inconsistencies, if any, noted in this material. This process had just begun and the fact that all these documents were in the Russian language, requiring meticulous and accurate translation, added to the time factor in assimilating the information by the Indian side.

The whole process came to an abrupt stop when the technology transfer agreement was annulled and downgraded as a supply contract. Subsequently, the Russians became reticent in parting with any information or giving clarifications since their interest was to keep ISRO as a tied customer for the cryo stages to be supplied by them for sustaining the GSLV programme. This was essential for the survival of their industries in the adverse economic environment that prevailed subsequent to the break-up of the USSR.

Then onwards, the progress on the indigenous cryo engine was very slow, since any issue of material/fabrication non-conformance or a subsystem test anomaly had to be resolved by the Indian team based on their own analysis and judgment as well as the limited exposure they were getting during the interaction with the Russian team when preparing the Russian-supplied cryo stage for the GSLV launches at Sriharikota.

In fact, it is more difficult when you try to follow someone else’s design without having any inputs on the “know-why” and then make it work successfully. The Russians did not share with ISRO any information on further modifications/updates they were carrying out in their cryo engine/stage based on their experience of ground tests. They flatly refused to give clarifications on specific critical elements such as the liquid oxygen [LOX] metering valve and the booster pump.

After the failure of the GSLV-D3 flight, how did you ensure this time that the ignition of the engine sustained for 12 long minutes in the vacuum of space?

When we attempted GSLV-D3 with the indigenous cryo upper stage, there was a great deal of confidence that the stage would work in flight. After all, by then we had carried out more than 6,000 seconds of cumulative engine hot test spread over 25 to 30 starts on the ground. We had also successfully completed a full-duration stage-firing lasting 720 seconds on ground, which the Russians themselves could not accomplish.

Yes, before GSLV-D3, there was a lot of debate on the CUS thrust frame, which had some deviation. However, there was a good understanding of its implication and a consensus that there was adequate margin. However, the abrupt stoppage of the liquid hydrogen booster pump within two seconds in flight came as a big surprise.

The GSLV-D3 boost phase flight was nominal and the starting parameters for the cryo upper stage in flight in terms of pressures and temperatures were perfect. The engine started with the tank head and the stored-gas-driven booster pump revving up to its designed revolutions per minute. The main igniter and the vernier engine igniters fired and there was ignition in the thrust chamber as could be detected by the acceleration spike. However, the engine start process did not progress due to the stoppage of the liquid hydrogen booster pump, which led to cavitation at the main turbo pump, and the propellant supply to the gas generator did not pick up as expected, leading to the abort of the entire start process.

Post GSLV-D3, from the flight data we could reconstruct the sequence of events leading to the failure. The abrupt stoppage of the liquid hydrogen booster pump was identified as the single event that caused this unsuccessful engine start in flight. All possible failure modes of the booster pump system were addressed and the hardware design as well as the manufacturing, assembly and testing processes were strengthened to avoid any of these failure modes.

This systematic approach proved right and the booster pump performance in the GSLV-D5 stage was close to nominal, leading to the successful start and full burn of the cryo stage.

As ISRO Chairman K. Radhakrishnan mentioned, you “took charge” after the failure of the April 2010 flight. How did you bring about this success? The morale must have been down, especially after the December 2010 failure with the Russian cryogenic stage as well.

It so happened that I took charge as Director, LPSC, at Valiamala on June 1, 2010, just after the GSLV-D3 failure. I was already chairing the GSLV-D3 Failure Analysis Committee [FAC], with a mandate to decipher what exactly went wrong in the indigenous cryo stage. Even before the GSLV-D3 flight, I was closely associated with the CUSP [Cryogenic Upper Stage Project] reviews as co-chairman and had played a crucial role in getting the first flight stage cleared as acceptable for putting on the GSLV-D3 vehicle. So, I was fully familiar with the systems and the LPSC team involved in the CUS development.

Frankly, the GSLV-F06 failure, where the Russian cryo stage was used, did not have any impact on the morale of the LPSC team since it was established in a matter of days that the reason for the failure was poor workmanship in the Russian stage in terms of the separation plane connectors anchoring arrangement. If at all, it strengthened our resolve to perfect the Indian cryo stage at the earliest to sustain the GSLV programme. Yes, the pressure on the CUS team and the LPSC was tremendous and the entire ISRO was looking for a successful GSLV flight with the Indian cryo stage.

My first job was to make the team understand and accept that we were working in a complex and challenging technology area where failures were not uncommon and that we should not be put off by these setbacks.

I recall the statement I made in my first interaction with the LPSC house journalPropulsion Today, where I said that propulsion systems development requires great amounts of grit and determination and it is not for the chicken-hearted.

I believe that propulsion systems are the most challenging to master and require strenuous efforts and a lot of field work under trying and risky conditions. Unfortunately, this is not fully comprehended by all.

I tried my best to make everyone appreciate the complexity of the task the LPSC team was handling and also put in the right perspective the vital role played by the LPSC in all ISRO programmes.

My basic training as a propulsion engineer and my empathy with the entire team cutting across ranks perhaps motivated them to give their best.

I enjoyed the trust and goodwill of each member of the LPSC family and I always did my best to highlight the dedicated efforts made by them in mastering this difficult technology. I am happy they didn’t let me down.

The April 15, 2010 flight, which was the first to use India’s own cryogenic stage, was to put a GSAT that weighed 2,220 kg into orbit. But GSLV-D5 used a 1,983-kg GSAT-14. So, the payload was lighter by 240 kg in the successful GSLV-D5 mission. Does this detract from its success?

After we faced the failure of GSLV-D3 and lost the GSAT-4 satellite, we thought it was prudent to play it safe and keep the primary mission objective of GSLV-D5 as proving the indigenous cryo stage in flight with the deployment of a satellite as an additional bonus. As such, we planned a lighter GSAT-14 with fewer transponders to keep the cost low. Also, the urge to make the CUS more robust in this second attempt, especially after the GSLV-F06 failure, attributed to the deflection and breaking of the shroud in the cryo stage, did indeed increase the inert weight of the CUS in GSLV-D5. All these factors made us attempt GSLV-D5 with GSAT-14 weighing less than GSAT-4 of GSLV-D3. However, this in no way brings down the significance of proving successfully the indigenous cryo stage. The restoration of payload to two-tonne-plus will happen in the very next flight through the optimisation of inert mass and a realistic pruning of margins.

I can confidently say that once we have a working cryo stage and the GSLV, enhancing its performance or payload incrementally is bound to happen in subsequent flights, as demonstrated in the case of the PSLV.

As you and SDSC (Satish Dhawan Space Centre) Director M.Y.S. Prasad mentioned, you made a spectacular comeback within four months of the scrubbing of the GSLV-D5 flight in August 2013 because the liquid fuel in the rocket’s second stage leaked. Did the Afnor tank develop cracks? Why was it not noticed during tests? Was quality assurance given a go-by in such an important mission?

The failure of the second stage [GS2] tank during the final phases of the countdown in August 2013 was due to inherent material characteristics of the AFNOR 7020 aluminium alloy, which is more prone to the stress corrosion cracking phenomenon, which manifests when stress is raised in a corrosive environment like propellant wetting. The crack in the material develops and grows very fast under these conditions and this is what precisely happened to the GS2 fuel tank when the pressure was increased to pre-launch level around T minus two hours.

This tank had successfully undergone all the inspection processes and also had been subjected to two proof pressure tests, the last one in April 2013. As such, this failure was purely due to inherent material proneness to cracking. In fact, AFNOR 7020 was being phased out and replaced by AA 2219.

For the PSLV second stage PS-2, this change was already implemented while the L40 stages [strap-on booster motors], which were developed at a later date for the GSLV, we started with AA 2219. However, for GS2, where the hardware off-take was slow due to the low launch rate of the GSLV, we still had the 7020 tanks and the AA 2219 tank was just getting ready. This was the reason for the use of the 7020 tank for the GSLV-D5 in the August 2013 launch. However, in the GS-2 stage assembled afresh for GSLV-D5 launch in January 2014, the AA 2219 material tank was incorporated.



You plan to do a suborbital flight of GSLV-Mk III in March this year without firing the cryogenic engine. What is a suborbital flight? What is its purpose? Is this flight going to carry a model of the crew module required for ISRO’s human space flight programme?

The suborbital flight test of GSLV Mk III, named LVM3-X mission, is primarily to characterise the new 600-tonne heavy-lift vehicle with two large solid strap-on boosters during its flight through the atmosphere.

The vehicle configuration will be in full and final form. However, the cryogenic stage, C25, will not be functional and will not develop any thrust. With this constraint, the vehicle can reach only about 5 km per second velocity, which falls short of the orbital velocity required. Hence the mission is suborbital, with the upper stage and payload re-entering atmosphere and falling back to the earth.

The whole objective of the LVM3-X mission is to validate a host of important parameters and characteristics of this totally new vehicle of relatively larger dimensions and large strap-ons. In effect, with this flight experience, when we attempt the first developmental launch, LVM3-D1, with a functional cryo stage, C25, we will have greater confidence that C25 will get an opportunity to ignite and perform in flight, which is indeed essential considering the enormous effort we put into realising the cryogenic engine and stage.

‘Will be able to repeat the success’ | Frontline

 

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@Developereo @ptldM3

 

[kurup]

You should have started this thread in Indian section .

 

[Developereo]

@Developereo @ptldM3

Far too long to read...
Anyways, good achievement.

 

[IND151]

Far too long to read...
Anyways, good achievement.

I have bolded important parts.