VISVESVARAYA TECHNOLOGICAL UNIVERSITYBELAGAVI-590 018

SEMINAR ON

CRYOGENIC ROCKET ENGINE

Bachelor Of EngineeringIn

MECHANICAL ENGINEERINGSubmitted By:

JAISON CYRIL (1SP13ME033)

Department of Mechanical EngineeringS.E.A COLLEGE OF ENGINEERING AND

TECHNOLOGYBENGALURU-560049

Cryogenic rocket engine

CONTENTS

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

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

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.

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

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

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

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

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.

ROCKET ENGINE POWER CYCLEGas pressure fed systemGas generator cycleStaged combustion cyclePump fed engine

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)

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.

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

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

THE FOUR PHASE OF COMBINATION IN THE THRUST CHAMBER ARE :-

Primary Ignition Flame Propagation Flame Lift off Flame Anchoring

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.

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.

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.

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.

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.

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

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.