NASA SLS - The new generation Rocket technology

1. NASA’S SPACE LAUNCH SYSTEM:
A heavy-lift platform for entirely new missions
Gokul Lakshmanan
M . Tech Thermal and Fluid Engineering

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

4. 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

5. 3 Proposed variations

6. specifications
Size
Diameter 8.4 m (core stage)
Stages 2
Capacity
Payload to
LEO
70,000 to 130,000 kg

7. 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

9. 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.

10. 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;

11. 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

12. Rearranging equation
Recalling that
or

13. So the final relation becomes

14. 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

15. 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

16. 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

18. •

19. Combustion Zones in Thrust Chamber
1. Injection/Atomization Zone
2. Rapid Combustion Zone
3. Stream Tube Combustion Zone

20. 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

21. 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

22. 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

23. 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

24. 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

25. 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

27. Electrical Bus system

28. 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

30. 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.

31. 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%

32. 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

33. • 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

34. • 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

35. 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

36. 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

37. • 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

38. 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

39. 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

40. 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

41. 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

42. • 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.

43. 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.

44. 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

45. 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

46. 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.

47. 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.

48. 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.

49. 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.