SCE-200 semi-cryo engine info

[K210]

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[antriksh]

Thrust (vacuum) - 2000 kN
Isp (vacuum) - 3285 N-s/kg
Chamber Pressure - 18 MPa
Mixture Ratio - 2.65
Thrust Throttling - 65-105 (% of nominal thrust)
Engine gimbal - 8 degrees (in two planes)

ISRO has started building thrust chamber test facility for semi-cryo engine

[antriksh]

Thrust (vacuum) - 2000 kN
Isp (vacuum) - 3285 N-s/kg
Chamber Pressure - 18 MPa
Mixture Ratio - 2.65
Thrust Throttling - 65-105 (% of nominal thrust)
Engine gimbal - 8 degrees (in two planes)

ISRO has started building thrust chamber test facility for semi-cryo engine

dosen't ISRO already have a test facility for testing liquid engines?

Not for a 2000 kN class semo-cryo engine

[kanaka]

Thrust (vacuum) - 2000 kN
Isp (vacuum) - 3285 N-s/kg
Chamber Pressure - 18 MPa
Mixture Ratio - 2.65
Thrust Throttling - 65-105 (% of nominal thrust)
Engine gimbal - 8 degrees (in two planes)

ISRO has started building thrust chamber test facility for semi-cryo engine

dosen't ISRO already have a test facility for testing liquid engines?

Not for a 2000 kN class semo-cryo engine

Need to send to RUSSIA for getting it tested?

[antriksh]

Thrust (vacuum) - 2000 kN
Isp (vacuum) - 3285 N-s/kg
Chamber Pressure - 18 MPa
Mixture Ratio - 2.65
Thrust Throttling - 65-105 (% of nominal thrust)
Engine gimbal - 8 degrees (in two planes)

ISRO has started building thrust chamber test facility for semi-cryo engine

dosen't ISRO already have a test facility for testing liquid engines?

Not for a 2000 kN class semo-cryo engine

Need to send to RUSSIA for getting it tested?

They are building test facilities in India.

[antriksh]

Rotary Vacuum Brazing Furnace

[cave_dweller]

A bit of history of SCE development in India.
http://www.frontline.in/static/html/fl2406/stories/20070406001404300.htm

[antriksh]

Single element Thrust chamber (experimental design) hot test

[antriksh]

Injector design experimental study for Oxidizer Rich Staged Combustion cycle

[antriksh]

Update:

1) Realisation of semi-cryogenic engine involves the development of performance-critical metallic and non-metallic materials and related processing technologies. 23 metallic materials and 6 non-metallic materials have been developed.

2) Characterisation of injector elements and hypergolic slug igniters with different proportion of Tri-ethyl Aluminium and Tri-ethyl Boron has been completed.

3) Sub-scale models of thrust chamber have been realised and ignition trials have been carried out successfully.

4) Single element thrust chamber hot test in stage combustion cycle mode was also conducted successfully.

Establishment of test facilities like Cold Flow Test Facility and Integrated Engine Test Facility are under various stages of realisation. Fabrication drawings are realised for all sub-systems and fabrication of booster turbo-pump and pre-burner subsystem commenced.

[vineethgk]

A pretty noob question that I have been having for some time in my mind...

Considering that LOX/Kerosene engines have been present from the dawn of the Space Age (R-7 comes to my mind), what exactly would be the primary engineering difficulty in designing a modern high thrust LOX/Kerosene engine like SCE-200? Would it be things like the design of turbopumps with sufficient mass-flow, dealing with combustion instability in thrust chamber etc...  all on account of the huge amount of thrust we are dealing with? And maybe meeting the weight and Isp targets as well?

Are these the challenges that necessitated the Chinese Space Agency and ISRO to approach Russians and Ukrainians respectively to get their LOX/Kerosene engine designs, despite having developed LOX/LH2 engines earlier?

[baldusi]

A pretty noob question that I have been having for some time in my mind...

Considering that LOX/Kerosene engines have been present from the dawn of the Space Age (R-7 comes to my mind), what exactly would be the primary engineering difficulty in designing a modern high thrust LOX/Kerosene engine like SCE-200? Would it be things like the design of turbopumps with sufficient mass-flow, dealing with combustion instability in thrust chamber etc...  all on account of the huge amount of thrust we are dealing with? And maybe meeting the weight and Isp targets as well?

Are these the challenges that necessitated the Chinese Space Agency and ISRO to approach Russians and Ukrainians respectively to get their LOX/Kerosene engine designs, despite having developed LOX/LH2 engines earlier?

Most initial engines were gas generators run fuel rich. But high performance requires staged combustion cycle. Which can only be done with oxygen rich preburners. That's close to an oxy torch environment. Developing the necessary metallurgy so it doesn't corrodes is the main obstacle. Then, the preburner needs to have something like 20 times the mass flow of a gas generator, so you get combustions instability problems on the preburner. And the main combustion chamber and all piping run at least at twice the pressure. And you really can't separate the preburner stability from the main combustion chamber, so you don't know if each part works until you fire the whole thing. For the rest, easy picy :-p

[vineethgk]

Most initial engines were gas generators run fuel rich. But high performance requires staged combustion cycle. Which can only be done with oxygen rich preburners. That's close to an oxy torch environment. Developing the necessary metallurgy so it doesn't corrodes is the main obstacle. Then, the preburner needs to have something like 20 times the mass flow of a gas generator, so you get combustions instability problems on the preburner. And the main combustion chamber and all piping run at least at twice the pressure. And you really can't separate the preburner stability from the main combustion chamber, so you don't know if each part works until you fire the whole thing. For the rest, easy picy :-p

Thanks for the explanation! Till now I was under the impression that combustion instability is something that affects the main combustion chamber alone. 

Any guesses as to what extent ISRO's prior experience of having built a Staged Combustion LOX/LH2 stage, albeit of much lower thrust, help in tackling these challenges?

[AJA]

Any guesses as to what extent ISRO's prior experience of having built a Staged Combustion LOX/LH2 stage, albeit of much lower thrust, help in tackling these challenges?

Guessing here. If you're talking about the CUS-7.5, they "built it" by reverse engineering a Russian engine for the full-cryo, staged combustion engine. They didn't start from basic principles, and therefore there isn't that experience with the fundamental research around cryogens. Not nearly as much as what the Russians and the Americans have.

Watch to get an idea of how long it took the Soviets. Also, starting at 40:21, you get an idea of the ridiculous environment in ox-rich pre-burners.

However, reverse engineering isn't trivial. They'd have learnt to handle cryogens on the ground, and design fuel flow systems without the cryogens boiling off in the middle of the line and causing "geysering". They'd have learnt to design the turbopumps to function without cavitation. They'd have learnt what materials they'd need to use to seal chambers, and valves and piping, and the metallurgy of the tanks and pipes, and combustion chambers to hold cryogens. They'd have built the closed loop electronic systems to monitor engine health, as well as instrumentation that can handle cryogenic temperatures (fuel level sensors etc.). They'd have built some electronic control unit for engine throttling - by altering mixture ratios and/or fuel flow to the pre-burner and/or main combustion chamber. A lot of that should be transferable.

However, the SCE-200 is a bigger engine. So they'd definitely have do revisit thermal and fluid flow solutions. Combustion instabilities, and engine characteristics would be different... so they'd have to start from scratch there. It's not staged combustion though, so that'd probably make things somewhat easier. Nozzle cooling would also have to be carried out with a different fluid (I'm assuming they used LH2 in the CUS-7.5), so that'd need attention. They might need a whole different injector/valve/piping design - in terms of materials used, as well as geometry, since kerosene has different physical and chemical properties. Finally, the entire thermodynamic cycle has changed.

As for why they didn't stick with the staged combustion design - that's been discussed here.

And you really can't separate the preburner stability from the main combustion chamber, so you don't know if each part works until you fire the whole thing.

Can't you test the turbopumps, the valves, the fuel flow, and the pre-burner - each individually? The only thing you'd need a complete engine for would be to test the combustion chamber. You'd need O/F rich hot gas and the remnant of F/O coming in to the combustion chamber. To get those, you'd have had to build a pre-burner .... so you've ended up building an engine. However, you could have really long piping and additional bleed/feed lines on a test rig (exploiting the fact that you're on the ground) - and thereby decouple the instabilities in the pre-burner, from those in the combustion chamber couldn't you?

[vineethgk]

Thanks AJA for the info and especially the YouTube link..

However, the SCE-200 is a bigger engine. So they'd definitely have do revisit thermal and fluid flow solutions. Combustion instabilities, and engine characteristics would be different... so they'd have to start from scratch there. It's not staged combustion though, so that'd probably make things somewhat easier.

Did you mean CE-20 and not SCE-200 here? I believe SCE-200 is a staged combustion engine as mentioned in the OP?

Btw, I read sometime back that ISRO got the SCE-200 design from Yuzhnoye OKB. Is this the Yuzhnoye counterpart we are talking about - The RD-810?
http://www.yuzhnoye.com/en/technique/rocket-engines/marching/rd-810/

If it is a design that is still under development by Yuzhnoye as mentioned in the link, is SCE-200/RD-810 some kind of ongoing joint ISRO-Yuzhnoye development, or did it stop with the sharing of design?

[baldusi]

CUS-7.5 is run fuel rich. H2 engines run better when fuel rich, since the thermal properties of H2 give a lot more power on the turbine. And don't have the metallurgy issues of O2. Regrettably, RP-1 would polymerize if run at preburner temperatures.
It's interesting to point out that the CUS-7.5 is air startable, but not restartable. But when they needed a higher thrust restartable engine they went with the easier gas generator cycle. That might say something about their development experience with staged combustion H2/LOX engine.

[vyoma]

http://www.newindianexpress.com/states/kerala/ISRO%E2%80%99s-Baby-Steps-at-Cheaper-Missions/2014/11/08/article2513064.ece

Officials at the LPSC HQ, Valiyamala, said they hoped to run the first major test in connection with the ‘semi-cryogenic’ engine project by November-end. What is special about the engine is that it uses kerosene as fuel instead of Liquid Hydrogen (LH2), the propellant used in cryogenic engines.

“This will be the first sub-system level test and we will be testing the booster pump for the oxidiser used in the engine,’’ LPSC director K Sivan said on Friday.  In both cryogenic and the semi-cryogenic engines, Liquid Oxygen is used as oxidiser, which helps the fuel to burn. In addition to being a low-cost technology, the use of highly refined kerosene (RP-1) will enable easier storage and handling.

The cold flow test facility at the LPSC unit in Mahendragiri, Tamil Nadu, where the test is to be conducted, is expected to be completed shortly, Sivan said. In fact, LPSC has had to postpone the test to November owing to the delay in its completion. An integrated test facility also is planned at Mahendragiri where the ‘hot test’ of the semi-cryo engine - in a hot test, the engine is fired - will be performed.

[abhishek]

Godrej Aerospace to make semi-cryogenic engines

The SME project was approved by the Government of India in January 2009 at a sanctioned cost of ₹1,798 crore. Department of Space’s Outcome Budget for 2014-15 says that the project is “in the initial stages”.

It expects the engine to be fully developed “after six years”.

Till the end of March 2013, ISRO had spent ₹155 crore on the project. Godrej will make six engines for ISRO. Vaidya said the company had begun work on three.

The SME is meant to power the future GSLV Mk III rockets as well as the heavy-life Unified Launch Vehicles, or ULV, which is today only a concept. The ULV will be a modular vehicle where the number of engines used will be based on the weight of the satellite or spacecraft.

The rocket will feature a combination of SME and an Indian cryogenic engine.

http://www.thehindubusinessline.com/news/has-been-supplying-the-vikas-engines-for-isros-rockets/article6705014.ece

[abhishek]

Meanwhile, ISRO has started forming concepts to develop a rocket that can put a 10-tonne satellite into orbit. This vehicle would require powerful engines. One candidate is the semi-cryogenic engine, using kerosene and liquid oxygen, whose design is now over. The hardware is being built and facilities being created. When ready, it will be an efficient lower stage with a thrust of 200 tonnes and controllable in flight, good enough to go into the lower stages of a large rocket. ISRO's plans are to use it in the heavy lifter and the reusable launch vehicle.

http://economictimes.indiatimes.com/articleshow/46294413.cms?utm_source=contentofinterest&utm_medium=text&utm_campaign=cppst

[TrevorMonty]

Meanwhile, ISRO has started forming concepts to develop a rocket that can put a 10-tonne satellite into orbit. This vehicle would require powerful engines. One candidate is the semi-cryogenic engine, using kerosene and liquid oxygen, whose design is now over. The hardware is being built and facilities being created. When ready, it will be an efficient lower stage with a thrust of 200 tonnes and controllable in flight, good enough to go into the lower stages of a large rocket. ISRO's plans are to use it in the heavy lifter and the reusable launch vehicle.

http://economictimes.indiatimes.com/articleshow/46294413.cms?utm_source=contentofinterest&utm_medium=text&utm_campaign=cppst

This LV should be similar to Atlas V ie 2 x 450klb 1st stage RP1 engines plus cryogenic upper stage.