Ppt on

Cryogenic EngineMiss : Deshpande.p.p.

Mr : atnurkar.a.d.Mr : bandewar.g.g.

Mr : sarwade.s.d.

Liquid propellant rocket enginesGeneral characteristics

Liquid propellant rocket engines are mostly widely used rocket engines because of many advantages that liquid propellants have. The first rockets used solid propellants because of the simplicity of their construction (just a barrel with gunpowder), but such engines were difficult to control. Chemistry and physics of combustion were undeveloped, combustion was unpredictable and it was nearly impossible to control it. Liquid rocket engines (LRE) were very promising: their thrust could be controlled by dosing propellant flow ratio with valves. Although nowadays the techniques of solid propellants have enormously advanced, liquid propellants retain their importance for the rocketry.

The main advantages of liquid propellants:• high specific impulse;• high thrust;• high thrust to weight ratio of the rockets;• easy to control.However, they also have certain disadvantages:• complexity of the construction;• it is impossible to achieve very high specific impulse;• they are difficult to scale, complexity grows quickly with growing the thrust;• some propellants are highly toxic (hydrazine and its variants) or cryogenic (hydrogen).

The cryogenic Engine

What is mean cryogenic ?• The study of the production and behavior of

materials at very low temperature. (below -150 degree C, -238 degree F, 123 K)

• A person studies elements have been subjected to extremely cold temperatures is called a cryogenicist.

Introduction 1. Geosynchronous satellite launch vehicles use Cryogenic

LOX/LH2 engine for their upper stage due to the higher specific impulse compared to earth storable engine. Major elements of Cryogenic system are propellant tanks, feed system and combustion chamber

2. For high thrust engine with longer burn duration, turbopump feed system are used

Cryogenic Engine /Stage

The cryogenic stage mainly consist of Fuel (LH2) and oxidiser (LOX)

tanks turbopump feed system, gas generator and combustion

chamber. The schematic of a typical stage is shown in figure

The cryogenic engine cycles normally used are gas generator cycle, staged combustion cycle

and expander mode cycle. These cycles are shown in figure2.

Expander cycle Staged combustion GG cycle cycle

• Particular cycle is chosen based on consideration like engine thrust, duration, state of art available in the organization etc. In the gas generator cycle (GG Cycle) the LOX and LH2 turbo pump are driven at two different speeds. It is done by using either a single turbine mounted on LH2 pump shaft and with gear box for reduced speed for LOX pump. Other option is two independent turbines for LOX and LH2 pumps mounted in either series mode or parallel mode. GG cycle is simple, and leads to independent development of subsystem like gas generator, turbopumps and combustion chamber. In GG Cycle the gas generator is run on small quantity of LOX/ LH2 tapped from the pumps. In this cycle there is overall loss of Isp due to lower Isp delivered by the turbine gas.

In expander cycle (EPC) the turbine drive gas, Gaseous hydrogen is tapped from the regeneratively cooled passage. The warm Gaseous hydrogen (GH2) after driving the turbine enters the combustion chamber where it burns with the oxygen to produce necessary thrust. This cycle is comparable to (SCC) as far as Isp loss is concerned. This cycle has yet another advantage that the turbine runs on GH2 at subatmospheric temperature. However the cycle has limitations that the engine size/thrust is limited by the magnitude of the heat extraction by hydrogen from the regenerating cooling passages.

System Description

System Description ISRO has embarked on a plan for realizing a gas generator cycle based cryogenic engine in the 12 to 20 tonne thrust range. For this engine LOX/LH2 turbopumps were mounted in the series mode of operation. The LH2 turbopump which needs more power is driven first by its turbine and the residual energy is used to run the LOX 3 turbine. The flow schematic is shown in fig

There are many parameters which are to be optimized before the detailed system design is taken up. These parameters are turbopump speed, gas generator, temperature, pressure ratio across each turbine, seals PV (Pressure and velocity product) number, bearing mounting and DN (bearing Diameter x speed) number, materials, pump net positive suction head etc. Detailed descriptions of these parameters are beyond the scope of this paper. The turbopumps were configured based on above optimization criteria. Major specifications are given in Table

PUMP LH2 LOXSpeed (rpm) 38000 15000Flow Rate (kg/sec) 7.5 40.5Head Rise (m) 13567 772

Efficiency (%) 60.5 59.5

TURBINE LH2 LOXSpeed (rpm) 38000 15000Inlet pressure (kg/cm2)

43 5.4

Inlet temperature (0c)

850 640

Flow rate (kg/sec) 1.24 1.17Power (H.P) 1675 481Efficiency (%) 0.46 0.46

Liquid Hydrogen Turbo Pump

• The turbopump consist of a cavitating inducer upstream of centrifugal pump. Two stage pumps are used. The volute used is a double tongue volute to reduce radical loads. Sealing System consist of mechanical Turbopump bellow seal for the pump side and hot gas seal for the turbine side. The rotating assembly is supported on two angular contact ball bearing cooled by liquid hydrogen. An additional bearing is provided upstream of the inducer to take care of pump over hang. In order to reduce the axial load on the bearing the balance piston mechanism is used. Pump is driven by a two row velocity compounded impulse turbine. Turbine is directly mounted on the pump shaft. Hot gas at 853 K received from LOX/LH2 gas generator runs the turbine. The exhaust form this turbine goes to LOX turbine inlet. Fig gives the LH2 turbopump assembly

Liquid Oxygen Turbo pump Liquid oxygen turbopump also consist of a cavitating inducer upstream of centrifugal pump. Volute is a double tongue volute. Bearing system consist of two angular contact ball bearings cooled by LOX. Third bearing is provided at the upstream of inducer for better support. The sealing system of this turbopump is very difficult. The difficulty comes due to availability of unburnt hydrogen in the turbine fluid. This unburnt hydrogen may lead to explosion if allowed to come in contact with the leaked LOX from pump. In order to avoid mixing of LOX and turbine hot gas, a helium purge seal is introduced between the bellow seal and hot gas seal. Helium gas at a pressure of 4bar is purged through this seal. This pressure gives the positive sealing between leaked pump fluid and leaked turbine fluid. LOX pump is driven by a two row velocity, compounded impulse turbine, which is directly mounted at the shaft. Hot gas after driving the LH2 pump drives the LOX pump.

Hardware and test facility realization All the elements of turbo pumps are realized in house as well utilizing external work centers. Turbine rotors in inconel and pump impeller in steel have been realized. These pumps and turbines were tested using water and GN2 as the working fluid respectively. The pump was driven by electric motor and turbine power was absorbed in dynamometer. This phase was completed under cold run testing of turbopump subsystem. Subsequent to the cold flow test turbopumps were tested individually under hot condition where pumps handled the actual fluid and turbine was driven by hot gas. Finally the turbopumps were tested in the integrated mode where LH2 and LOX turbo pump were driven in series mode of operation, pump handling respective propellent and turbine driver by gas generator hot product. The turbo pump performance for LH2 and LOX turbo pump is shown in fig 6 & 7.

Conclusion • The cryogenic turbo pump (LOX / LH2) which are most

complex system in the cryogenic propulsion has been developed successfully. This demonstrates the competence of ISRO in this area.

References

• ISRO• Turbomachines for Cryogenic Engines

N.K Gupta LPSC/ISRO Trivandrum

THANKS