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NASA SLS - The new generation Rocket technology

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  1. 1. NASA’S SPACE LAUNCH SYSTEM: A heavy-lift platform for entirely new missions Gokul Lakshmanan M . Tech Thermal and Fluid Engineering
  2. 2. INTRODUCTION • Heavy expendable launch vehicle • Designed by NASA • Replace NASA’s retired Space Shuttle • SLS launch vehicle is to be upgraded over time with more powerful versions
  3. 3. VARIATIONS Block 1 • Lifts a payload of 70 metric tons to LEO Block 1b • Lifts a payload of 105 metric tons to LEO Block 2 • Lifts a payload of more than 130 metric tons to LEO
  4. 4. 3 Proposed variations
  5. 5. specifications Size Diameter 8.4 m (core stage) Stages 2 Capacity Payload to LEO 70,000 to 130,000 kg
  6. 6. 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
  7. 7. 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 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.
  8. 8. 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;
  9. 9. 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
  10. 10. Rearranging equation Recalling that or
  11. 11. So the final relation becomes
  12. 12. 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
  13. 13. 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
  14. 14. 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. • 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
  15. 15.
  16. 16. Combustion Zones in Thrust Chamber 1. Injection/Atomization Zone 2. Rapid Combustion Zone 3. Stream Tube Combustion Zone
  17. 17. 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. Chemical reactions occur in this zone, but the rate of heat generation is relatively low
  18. 18. 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 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
  19. 19. 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
  20. 20. 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
  21. 21. Solid Rocket Booster (SRB) • 2 Solid fuel rocket boosters used for primary propulsion • Provided the majority of the thrust during the first two minutes of flight. • Thrust :16000 kN • Burn time 124 seconds
  22. 22. Components 1. Hold-down posts • Each solid rocket booster has four hold-down posts • Hold-down bolts hold the SRB and launcher platform together • Hold down nuts contains NASA Standard Detonators(NSD) which were ignited at SRB ignition commands • NSD ignite and splits the nut into two or more parts • Hold-down bolt travels downward • The SRB bolt is 710 mm long and 89 mm in diameter
  23. 23. Electrical Bus system
  24. 24. 3. Hydraulic power units • Two independent Hydraulic Power Units (HPUs) on each SRB • The gas generator decompose the hydrazine (fuel) into hot, high-pressure gas • A turbine converted this into mechanical power, driving a gearbox. • The gearbox drive the fuel pump hydraulic pump. • The hydraulic pump speed was 3600 rpm and supplied hydraulic pressure of 21.03 ± 0.34 Mpa • Hydraulic pressure is used to drive Thrust vector controller
  25. 25. 4. Thrust vector control • to move the nozzle up/down and side-to-side. • This provided thrust vectoring to help control the vehicle in all three axes (roll, pitch, and yaw). • Each SRB servo actuator consist of four independent, servo valves that receive signals from the drivers. • Each servo valve control one actuator ram and thus nozzle to control the direction of thrust.
  26. 26. 5. Propellant Component Description % by weight Ammonium Perchlorate (NH4ClO4) oxidizer 69.6% Aluminum fuel 16% Iron oxide a catalyst 0.4% Poly butadiene acrylonitrile Serves as a binder that hold the mixture together and acted as secondary fuel 12.04%
  27. 27. Operation sequence 1. Ignition • Ignition can occur only when a manual lock pin from each SRB has been removed. • The ground crew removes the pin during prelaunch activities at T minus five minutes • The solid rocket booster ignition commands are issued when four cryogenic engines are at or above 90% rated thrust
  28. 28. • The fire commands cause the NSDs on the SRB to fire. • The booster charge ignites the propellant in the SRB which fires down the entire vertical length
  29. 29. • This ignites the solid rocket propellant along its entire surface area instantaneously. • At t minus zero, the two SRBs are ignited, under command of the four onboard computers
  30. 30. 2. Separation • The SRBs are jettisoned from SLS at altitude, about 45 km. • SRB separation is initiated when chamber pressure of both SRBs is less than or equal to 340 kPa. • The SRBs separate from the SLS within 30 milliseconds of the firing command. • Attachment point consists a nut-bolt system • Detonating the NSD via electrical system separates the SRB’s
  31. 31. 3. Descent and Recovery • The SRBs are jettisoned from the SLS at 2 minutes and an altitude of about 45 km. • After continuing to rise to about 67 km the SRBs begin to fall back to earth • Once back in the atmosphere are slowed by a parachute system to prevent damage on ocean impact
  32. 32. • Nose cap separation occurs at a nominal altitude of 5km, about 218 seconds after SRB separation. • This triggers the parachute to open and SRB falls to ocean • SRB is later recovered by US Navy
  33. 33. Orion Multi-Purpose Crew Vehicle • Carry a crew of up to four • Beyond low earth orbit • Currently under development by NASA • Sustain the crew during space travel • Provide safe re-entry from deep space. • Also provides an emergency launch abort capability
  34. 34. DIMENSIONS Height: 3.3 m Diameter: 5 m Pressurized volume: 19.56 m3 Capsule mass: 8,913 kg Service Module mass: 12,337 kg Total mass: 21,250 kg Service module propellant mass: 7,907 kg
  35. 35. Launch Abort System Provides crew escape during launch pad and ascent emergencies Service Module Power, propulsion and environmental control support to the Crew Module Crew Module Human habitat from launch through landing and recovery Spacecraft Adapter Orion-to-Space Launch System (SLS) structural interface
  36. 36. Crew module (CM) • Reusable • Provides a habitat for the crew, provides storage for consumables and research instruments • Only part of Orion MPCV that returns to earth after each mission • It will have more than 50% more volume than the Apollo capsule • Carry four to six astronauts
  37. 37. • The CM is constructed of aluminum-lithium alloy • The CM is covered with thermal protection system • Reusable recovery parachutes to slow down the decent of spacecraft into earth • Designers claim that the Orion is designed to be 10 times safer than the space shuttle.
  38. 38. Advanced technologies used in Orion crew module • Glass cockpit: Features digital displays, typically large LCD screens, compared to the traditional style of analog dials and gauges • An "Auto dock" feature : Allows the orion spacecraft to control itself automatically and dock with international space station in space • Improved waste-management facilities, with a miniature toilet • A nitrogen/oxygen mixed atmosphere at sea level 101.3 kPa pressure. • Much more advanced computers than on previous crew vehicles.
  39. 39. Thermal Protection System Includes various types of material covering the Orion for thermal protection • Reinforced carbon–carbon (RCC)- Used where reentry temperature exceeded 1260 °C. • High-temperature reusable surface insulation (HRSI) tiles made of coated Silica ceramics. Used where reentry temperature was below 1260 °C. • Fibrous refractory composite insulation (FRCI) tiles or Alumina-borosilicate fiber, used to provide improved strength, durability, resistance to coating cracking and weight reduction. • Flexible Ceramic Insulation Blankets (FCIB), flexible blanket-like surface insulation. Used where reentry temperature was below 649 °C. • Felt reusable surface insulation (FRSI). Used where temperatures stayed below 371 °C. Its a low grade FRCI
  40. 40. Radiation Shielding • Materials rich in hydrogen and carbon are known to be effective shielding materials • Usually lead coating is used in the spacecraft • Water is also known to be an effective shielding material
  41. 41. Orion uses various techniques to protect the astronauts from space radiations • Lead coating • Aluminum coating and aluminum foil • Maximizing the amount of material that can be placed between the crew and the outside environment • Includes supplies, equipment, seats, as well as water and food.
  42. 42. Service Module Provides primary power and propulsion Functions i. Supports the crew module from launch through separation before reentry. ii. Provides in-space propulsion capability iii. Provides the water and oxygen needed for a habitable environment iv. Generates and stores electrical power v. Maintains the temperature of the vehicle's systems and components vi. Transport unpressurized cargo and scientific payloads.
  43. 43. Launch Abort System (LAS) • Crew safety system • Quickly separate the capsule from rocket in case of a launch abort emergency, such as an impending explosion. • The system is typically controlled by a combination of automatic rocket failure detector , and manual control by the crew commander.
  44. 44. i. Mounted above the capsule ii. Delivers a relatively large thrust for a brief period of time to send the capsule a safe distance away from the launch vehicle iii. The capsule's parachute recovery system can be used for a safe landing on ground or water.