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NASA SLS Cryogenic Engine - Complete Explanation


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4 cryogenic engines used in core stage by SLS.

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NASA SLS Cryogenic Engine - Complete Explanation

  2. 2. Contents • Cryogenic Engine • Construction • One-dimensional analysis of gas flow in rocket engine nozzles • Rocket cycles • Liquefaction and storage of cryogenic fuel • Combustion Zones in Thrust Chamber • Regenerative Cooling
  3. 3. Cryogenic Engine • 4 engines • Used in core stage • Initial flights will use engines left over from the Space Shuttle program • Later flights use cheaper version of the engine not intended for reuse • Use LH / LOX • LH at -2530c and LOX at -1830c • Provides 7440KN thrust
  4. 4. Construction • Combustion chamber • Pyrotechnic initiator (igniter): zirconium – potassium perchlorate mixture • Fuel injector and fuel pumps • Oxidizer pumps • Gas turbine • Fuel tanks • Rocket engine nozzle
  6. 6. ONE-DIMENSIONALANALYSIS OF GAS FLOW IN ROCKET ENGINE NOZZLES The analysis of gas flow through de Laval nozzles involves a number of assumptions: 1. The combustion gas is assumed to be an ideal gas. 2. The gas flow is isentropic i.e., at constant entropy, as the result of the assumption of non-viscous fluid, and adiabatic process. 3. The gas flow is constant during the period of the propellant burn. 4. The gas flow is non-turbulent 5. The flow behavior is compressible since the fluid is a gas.
  7. 7. When there is no external work and heat transfer, the energy equation becomes Differentiation of continuity equation, and dividing by the continuity equation We have;
  8. 8. For isentropic process ds=0 and combining equations Differentiation of the equation and dividing the results by the equation Obtaining an expression for dU/U from the mass balance equation
  9. 9. Rearranging equation Recalling that or
  10. 10. So the final relation becomes
  11. 11. Flow characteristics • At entry into the nozzle flow is subsonic • As the throat constricts, the gas is forced to accelerate until at the nozzle throat • At throat velocity becomes sonic • From the throat the cross-sectional area increases • The gas expands and the linear velocity becomes progressively more supersonic.
  12. 12. The linear velocity of the exiting exhaust gases can be calculated using the following equation M = the gas molecular mass, kg/kmol ɣ = cp/cv (isentropic expansion factor) Pe = absolute pressure of exhaust gas at nozzle exit p = absolute pressure of inlet gas
  13. 13. Types of Expansion 1. Under expanded 2. Ambient 3. Over expanded 4. Grossly over expanded Expansion varies with altitude If the exit pressure is too low jet separate from the nozzle. This is often unstable, and the jet will generally cause large off-axis thrusts and may mechanically damage the nozzle.
  14. 14. Specific Impulse The thrust of a rocket engine nozzle can be defined as Or;
  15. 15. The specific impulse Isp is the ratio of the of thrust produced to the weight flow of the propellants.
  16. 16. Rocket cycles • Pressure-fed engine cycle • Expander cycle • Gas-generator cycle • Staged combustion cycle
  17. 17. Pressure-fed engine • Propellant tanks are pressurized to supply fuel and oxidizer to the engine, eliminating the need for turbo pumps. • Pressurized Helium is often used
  18. 18. Expander cycle • Heat from the nozzle and combustion chamber powers the fuel and oxidizer pumps. • Uses a gas generator of some kind to start the turbine and run the engine until the heat input from the thrust chamber and nozzle skirt in sufficient enough to run the turbine
  19. 19. Gas-generator cycle • Some of the fuel and oxidizer is burned separately to power the pumps and then discarded. Most gas-generator engines use the fuel for nozzle cooling. • Some of the fuel used to cool the nozzle and combustion chamber.
  20. 20. Staged combustion cycle • All of the fuel and a portion of the oxidizer are fed through the pre-burner, generating fuel-rich gas. After being run through a turbine the gas is injected into the combustion chamber and burned. • Advantage: No loss of heat compared to gas generator cycle • USED IN SPACE LAUNCH SYSTEM
  21. 21. How to Liquefy cryogenic fuel Critical temperature for hydrogen -2530c 1. At first gaseous hydrogen is compressed to 180 atm. 2. Compressed gas is cooled by allowing it to expand rapidly. 3. The cooled expanded gas then passes through a heat exchanger where it cools the incoming compressed gas 4. The cycle is repeated until hydrogen liquefies
  22. 22. How to store cryogenic fuel • Cryogenic Dewar wall: Vacuum flask used for storing cryogenic fuels • Have walls constructed in two or more layers of silver with a high vacuum maintained between the layers. • Provides very good thermal insulation • Reduces the rate at which the contents boils away • Dewar allow the gas to escape through an open top to avoid risk of explosion • More sophisticated Dewar trap the gas above the liquid, and hold it at high pressure • This increases the boiling point of the liquid, allowing it to be stored for extended periods
  23. 23. The cryogenic Engine • Fuel and oxidizer from the rocket's core stage will flow directly into the fuel lines • The fuel and oxidizer each branch out into separate paths to each engines • In the engine fuel and oxidizer follows various paths to reach into the combustion chamber
  24. 24. Fuel Injection • Injector introduce and meter the flow of liquid propellants to the combustion chamber • Hole pattern of the injector is closely related to feed passages within the injector • A large manifold allows low passage velocities and good distribution. • A small manifold allows for a lighter weight injector and reduces the amount of "dribble" flow after the main valves are shut.
  25. 25. • Dribbling results in afterburning, which is an inefficient irregular combustion that gives a little "cutoff" thrust after valve closing. • For accurate vehicle velocity requirements, the cutoff impulse has to be very small. Types 1. Impinging-stream-type 2. Non-impinging 3. Coaxial injector
  26. 26. 1. Impinging-stream-type : Fuel and oxidizer streams impinge upon each other. Impingement aids atomization of the liquids into droplets also its distribution. 2. Non-impinging type: relies on turbulence and diffusion to achieve mixing
  27. 27. 3. Coaxial injector It has been used for liquid oxygen and gaseous hydrogen. This type of injector is used by SLS
  28. 28. Combustion Zones in Thrust Chamber 1. Injection/Atomization Zone 2. Rapid Combustion Zone 3. Stream Tube Combustion Zone
  29. 29. Injection/Atomization Zone 1. Injection, atomization and vaporization occurs here 2. Fuel and Oxidizing agent are introduced in this zone at velocities between 7 and 60 m/sec 3. The individual jets break up into droplets by impingement of one jet with another 4. Heat is transferred to the droplets by radiation from rapid combustion zone and by convection from moderately hot gases in the first zone. 5. Chemical reactions occur in this zone, but the rate of heat generation is relatively low
  30. 30. • If one of the propellants is a gas: • this occurs with liquid oxygen and gaseous hydrogen propellant from thrust chambers or pre-combustion chambers • The gas usually has a much higher injection velocity (above 120 m/sec) than the liquid propellant • This cause shear forces to be imposed on the liquid jets • Thus Coaxial injectors are used in such cases
  31. 31. Rapid Combustion Zone 1. Intensive and rapid chemical reactions occur at increasingly higher temperature 2. The mixing is aided by local turbulence and diffusion of the gas species 3. The rate of heat release increases greatly and this causes the specific volume of the gas mixture to increase 4. Axial velocity increase by a factor of 100 or more. 5. Gas flows from hot sites to colder sites. 6. Rapid fluctuations in pressure, temperature, density and radiation emissions occurs
  32. 32. Stream Tube Combustion Zone 1. Axial velocities are high (200 to 600 m/sec) 2. Streamlines are formed and there is relatively little turbulence 3. Residence time in this zone is very short 4. Usually less than 10 milliseconds 5. Volumetric heat release being approximately 370 MJ/m3sec 6. The higher temperature in the chamber causes chemical reaction rates to be several times faster
  33. 33. Regenerative Cooling • It is a configuration in which some or all of the propellant is passed through tubes around the nozzle to cool the engine • The heated propellant is then fed into a special gas generator or injected directly into the main combustion chamber • This is done because the nozzle material cannot withstand the heat produced by combustion • So the fuel itself is used to take away the heat