CRYOGENIC ROCKET ENGINE PPT

1. VISVESVARAYA TECHNOLOGICAL UNIVERSITY
BELAGAVI-590 018
SEMINAR
ON
CRYOGENIC ROCKET
ENGINE
Bachelor Of Engineering
In
MECHANICAL ENGINEERING
Submitted By:
JAISON CYRIL (1SP13ME033)
Department of Mechanical Engineering
S.E.A COLLEGE OF ENGINEERING AND
TECHNOLOGY
BENGALURU-560049

2. Cryogenic
rocket engine

3. CONTENTS
1 CRYOGENIC ?
2 INTRODUCTION
3 HISTORY OF CRYOGENIC
4 CONSTRUCTION
5 WORKING PRINCIPLE
6 APPLICATIONS
7 ADVANTAGES
8 DISADVANTAGES
9 CONCLUSION

4. CRYOGENIC
 Cryogenics is derived from Greek, Kryos means cold, genes
means production.
 Cryogenic is the study of production and behaviour of material at
very low temperature. (below
-150 ˚C, 123 K, -238 ˚F)
 Oxygen liquefies at -183 ˚C
 Hydrogen liquefies at -253 ˚C

5. INTRODUCTION
 A cryogenic rocket engine is a engine which use
cryogenic fuel.
 Cryogenic fuel are fuel that requires storage at
extremely low temperature in order to maintain them
in a liquid state.
 Various cryogenic fuel-oxidizer combination have
been fired but the combination of liquid hydrogen
(LH2), and the liquid oxygen (LOX) oxidizer is mostly
used.

6.  The RL10 was the first liquid hydrogen
cryogenic rocket engine to be built in the
United States, and development of the
engine by Marshall Space Flight
Center and Pratt & Whitney began in the
1950s, with the first flight occurring in
1961.
 These engines were one of the main
factors of NASA’s success in reaching the
Moon by the Saturn V rocket

7.  The specifications and key characteristics of the
engine are:
 Operating Cycle – Staged combustion
 Propellant Combination – LOX / LH2
 Maximum thrust (Vacuum) – 75 Kn
 Operating Thrust Range – 73.55 kN to 82 kN
 Chamber Pressure (Nom) – 58 bar
 Engine Mixture ratio (Oxidizer/Fuel by mass) –
5.05
 Engine Specific Impulse - 454 ± 3 seconds
(4.452 ± 0.029 km/s)
 Engine Burn Duration (Nom) – 720 seconds
 Propellant Mass – 12800 kg
C E 7.5

8. C E 20
 The specifications of the engine as listed on the:
 Operating Cycle - Gas Generator
 Propellant Combination - LOX / LH2
 Thrust Nominal (Vacuum) - 200 KN
 Operating Thrust Range - 180 KN to 220 KN (To be
set at any fix values)
 Chamber Pressure (Nom) - 6 Mpa
 Engine Mixture ratio (Oxidizer/Fuel by weight) - 5.05
 Engine Specific Impulse - 443 ± 3 seconds
(4.344 ± 0.029 km/s)
 Engine Burn Duration (Nom) - 595 seconds
 Total Flow rate - 462 kg/s
 Nozzle Area ratio – 100
 Mass - 588 kg

9. CONSTRUCTION
 The thrust chamber or combustion chamber
 pyrotechnic igniter
 fuel injector
 fuel turbo-pumps
 gas turbine
 cryo valves
 Regulators
 The fuel tanks
 rocket engine
 nozzle

10. WORKING PRINCIPLE
The basic principle driving a rocket engine are:-
 Newton third law of motion.
 Law of conservation of momentum.
 In principle, cryogenic rocket engine drives thrust like all
other rocket engine by accelerating an impulse carrier to
high speed.
 The chemical energy stored in the fuel is converted into
kinetic energy by burning fuel in the thrust chamber and
subsequently expansion in nozzle to produce thrust.

12. ROCKET ENGINE POWER CYCLE
Gas pressure fed system
Gas generator cycle
Staged combustion cycle
Pump fed engine

13. GAS PRESSURE FED
SYSTEM
 The pressure-fed engine is a class of
rocket engine.
 Helium is used as a pressurize the
propellant tank to force the fuel and
oxidizer to the combustion chamber.
 Tank pressure should exceed the
combustion chamber pressure
 It’s a simple plumbing and unreliable
turbopumps .
 If the fuel is hypergolic then they burn
as contact, if not igniter burner is
required to ignite.
 Usage:-( quad rocket, Aquarius launch
vehicles)

14. GAS GENERATOR
CYCLE
 The gas generator is a power cycle of a
bipropellant rocket engine.
 Some of the propellant is burned in a
gas generator and resulting hot gas is
used to power the engine’s pumps.
 Some of the fuel in a gas generator
cycle may be used to cool the nozzle
and combustion chamber.
 Without any rocket combustion
chamber and nozzle heat mitigation,
the engine would fail catastrophically.
 usage:-(F-1 rocket engine , vulcain, CE
20). etc.

15. STAGED COMBUSTION
CYCLE
 It is a bipropellant rocket engine
 One propellant is sent through preburner
and partially burned using a small portion of
second propellant
 The resulting hot gas is used to power
engine turbine and pumps, then injected
into main combustion along with the
remainder of second propellant to complete
combustion
 In staged combustion all the cycle of gases
and heat go through the combustion
chamber

16. PUMP FED ENGINE
 It is a power cycle of bio
propellant rocket engine
 It uses dual electrical pumps to
increase the pressure from the
tank to combustion chamber
 Pump is actuated by an
electrical motor , fed by a
battery bank
 Inverter convert DC current to
AC

17. THE FOUR PHASE OF COMBINATION IN
THE THRUST CHAMBER ARE :-
 Primary Ignition
 Flame Propagation
 Flame Lift off
 Flame Anchoring

18. PRIMARY IGNITION
 Begins at the time of deposition of the
energy into the shear layer and ends when
the flame front has reached the outer limit
of the shear layer.
 Starts interaction with the recirculation
zone.
 Phase typically lasts about half a
millisecond
 It is characterized by a slight but distinct
downstream movement of the flame.
 The flame velocity more or less depends
on the pre-mixedness of the shear layer
only.

19. FLAME PROPAGATION
 This phase corresponds to the time span for the
flame reaching the edge of the shear layer,
expands into in the recirculation zone and
propagates until it has consumed all the
premixed propellants
 This period lasts between 0.1 and 2 ms.
 It is characterized by an upstream movement of
the upstream flame front until it reaches a
minimum distance from the injector face plate.
 It is accompanied by a strong rise of the flame
intensity and by a peak in the combustion
chamber pressure.

20. FLAME LIFT OFF
 Phase starts when the upstream flame front
begins to move downstream away from the
injector because all premixed propellants in
the recirculation zone have been
consumed until it reaches a maximum
distance.
 This period lasts between 1 and 5ms.
 The emission of the flame is less intense
showing that the chemical activity has
decreased.
 The position where the movement of the
upstream flame front comes to an end, the
characteristic times of convection and
flame propagation are balanced.

21. FLAME ANCHORING
 This period lasts from 20ms to more than 50ms ,
depending on the injection condition.
 It begins when the flame starts to move a second
time upstream to injector face plate and ends when
the flame has reached stationary conditions.
 During this phase the flame propagates upstream
only in the shear layer.
 Same as flame lift-off phase the vaporization is
enhanced by the hot products which are entrained
into the shear layer through the recirculation zone.

22. ADVANTAGE
 High Energy per unit mass:
Propellants like oxygen and hydrogen in liquid form give very high amounts of
energy per unit mass due to which the amount of fuel to be carried aboard the
rockets decreases.
 Clean Fuels
Hydrogen and oxygen are extremely clean fuels. When they combine, they give
out only water. This water is thrown out of the nozzle in form of very hot vapour.
Thus the rocket is nothing but a high burning steam engine
 Economical
Use of oxygen and hydrogen as fuels is very economical, as liquid oxygen costs
less than gasoline.

23. DISADVANTAGE
 Boil off Rate
 Cryogenic fluid are difficult to store for longer period
 Highly reactive gases
 Zero gravity condition
 Leakage
 High density requires larger tank
 Hydrogen Embrittlement

24. THE NEXT GENERATION OF ROCKET
ENGINE
 All rocket engines burn their fuel to generate thrust . If any other engine
can generate enough thrust, that can also be used as a rocket engine
 There are a lot of plans for new engines that the NASA scientists are still
working with. One of them is the “ Xenon ion Engine”. This engine
accelerate ions or atomic particles to extremely high speeds to create
thrust more efficiently. NASA's Deep Space-1 spacecraft will be the first to
use ion engines for propulsion.
 There are some alternative solutions like Nuclear thermal rocket engines,
Solar thermal rockets, the electric rocket etc.