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Rapid Response Launcher System
(RRLS)
Ramirez, Uriel
Johnson, Lindani
Gonzalez, Austin
Solomon-Williams, Cordarryl M.
Smith, Joseph
Knill, Christian
Problem Statement
A rocket system to set a network of Search & Rescue Satellites?
“
“In light of the MH 370 tragedy in the spring of 2014,
an interest has arisen to investigate the feasibility of
a rapid response launcher system for a constellation
of simple, lightweight search satellites with a
minimum orbital lifetime of 6 months.”
Payload - 100lb S&R Satellite
Orbit - Circular Sun Synchronous at an altitude of 165 km
Launch Location - Poker Flat Research Range, AK
65.1167° N Lat, 147.4612° W Lon, 436.7 m above sea level
Mission Requirements
Extraneous Requirements
Environmentally Friendly Propellant System
Total system weight less than 10,000kg
Minimum orbital lifetime 6 months
Missing Planes
Poker Flat Research Range, AK
MH 370
Orbital Elements from STK
Period 5267.32 sec
SemiMajor Axis 6543.14 km
Eccencenity 9.33702 e-018
Inclination 96.2587 deg
Argument of Perigee 0 deg
True Anomaly 359.924 deg
Ascending Node 126.994 deg
Satellite Tool Kit Simulation Verification
(2D Graphic of Earth)
Satellite Tool Kit Simulation Verification
(3D Graphic of Earth)
Preliminary Design?
ΔV Design 1
Burnout Velocity -
Velocity of Launch site -
ΔV Design 2
Velocity Needed -
Design Velocity -
Assuming losses of 0.9km/s
Three stages:
Stage 1 will produce a ΔV = 3.5 km/s
Stage 2 will produce a ΔV = 3.5 km/s
Stage 3 will produce a ΔV = 2.25 km/s
Orbit Design
Source: Delta II Payload Planners Guide December 2006
Three Stage Design
Pros
◇ Reasonable ISP will
meet our needs
◇ Better finert values
Cons
◇ More complicated due
to more staging
◇ Larger inert mass
System Level
Performance?
Propellant Choice
Choosing H202 and
HTPB allowed us to
have a reasonable
ISP and good finert
that will not be
technologically
challenging
F2/H2 and O2/H2 had
a large ISP, but too
toxic.
O2/RP-1- Needed
oxygen tanks
Propellant Trade Offs
An initial mass of about 5500
kg, ISP of 285s and finert of
0.28 was chosen.
The ISP of 285 allowed us to
account for pressure losses in
our system with the chosen
propellant.
Preliminary Sizing: Stage 1
Helium pressurizes the propellant
tanks to force the fuel and
oxidizer tanks to the combustion
chamber
Pros:
◇ Simple, due to less
components
◇ Easy to maintain
Cons:
◇ Extra weight added because
of the pressurant tanks
Pressure-fed engine
Injector in a liquid rocket engine mixes the
fuel with the oxidizer to produce efficient
and stable combustion
This figure shows an injector designed with
propellant valves (remote control)
Injector
Injector Types
AUTOCAD DRAWING
Nozzle efficiency 97 %
Nozzle Sizes
Our design conditions were best met by
ablative cooling
The ablative material absorbs the heat as it
ablates
Cooling Type
Pros:
◇ Simple
◇ Capable of stopping
and restarting the
engine, as long there is
ablative material left
Cons:
◇ Increase of weight
◇ Limited life in the
engine (usually less
than 2000 seconds)
Typical ablative materials: Silica, Quartz,
or Carbon Cloth and resin composites
Cooling Type Trade offs
Material Density (kg/m3) Ultimate Tensile
Strength (GPa)
Specific Ultimate
Tensile Strength
(Gpa/(kg/m3))
2219 Aluminum 2800 0.413 15.04
Titanium 4460 1.23 28.81
4130 Steel 7830 0.892 11.23
Graphite 1550 0.895 58.88
SOURCE: Space Propulsion Analysis and Design, Humble, Henry, and Larson
Pressure Tank Material
Titanium has the best
mechanical properties,
but is difficult to work
with and extremely
expensive.
Steel is cheap and easy
to work with, but does
not have the properties
required for our design.
Composites meet the
property requirements
and are lightweight.
Our rocket cost
increases, within
reason.
Material Trade
offs
Due to the addition of the pressurant tanks, composite
materials were chosen to minimize the mass and
therefore increase the ΔV
This increased the cost of our rocket.
Final Predicted Design
Results?
ISP (s)
(from
RPA
code)
Initial
Mass (kg)
Propellant
Mass (kg)
Thrust
(kN)
Burn time
(s)
ΔV (m/s)
(Actual)
Stage 1 287 5550 3904 2698 4 3348
Stage 2 287 900 637 53 33 3317
Stage 3 287 200 109 12 26 2798
Total 9463
The actual ΔV, considering all the actual masses, is 2.5% bigger than
design.
Final Predicted Results
Altitude Simulation
Velocity Simulation
Mass Simulation
Conclusion
Theoretically, it is feasible.?
Justifications
1. Small propulsion
system to achieve an
orbit
2. Uses current
technologies proven to
work
3. Performs a
meaningful and
desirable task
THE END

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Rapid Response Launcher System

  • 1. Rapid Response Launcher System (RRLS)
  • 2. Ramirez, Uriel Johnson, Lindani Gonzalez, Austin Solomon-Williams, Cordarryl M. Smith, Joseph Knill, Christian
  • 3. Problem Statement A rocket system to set a network of Search & Rescue Satellites?
  • 4. “ “In light of the MH 370 tragedy in the spring of 2014, an interest has arisen to investigate the feasibility of a rapid response launcher system for a constellation of simple, lightweight search satellites with a minimum orbital lifetime of 6 months.”
  • 5. Payload - 100lb S&R Satellite Orbit - Circular Sun Synchronous at an altitude of 165 km Launch Location - Poker Flat Research Range, AK 65.1167° N Lat, 147.4612° W Lon, 436.7 m above sea level Mission Requirements
  • 6. Extraneous Requirements Environmentally Friendly Propellant System Total system weight less than 10,000kg Minimum orbital lifetime 6 months
  • 7. Missing Planes Poker Flat Research Range, AK MH 370
  • 8. Orbital Elements from STK Period 5267.32 sec SemiMajor Axis 6543.14 km Eccencenity 9.33702 e-018 Inclination 96.2587 deg Argument of Perigee 0 deg True Anomaly 359.924 deg Ascending Node 126.994 deg
  • 9. Satellite Tool Kit Simulation Verification (2D Graphic of Earth)
  • 10. Satellite Tool Kit Simulation Verification (3D Graphic of Earth)
  • 12. ΔV Design 1 Burnout Velocity - Velocity of Launch site -
  • 13. ΔV Design 2 Velocity Needed - Design Velocity - Assuming losses of 0.9km/s
  • 14. Three stages: Stage 1 will produce a ΔV = 3.5 km/s Stage 2 will produce a ΔV = 3.5 km/s Stage 3 will produce a ΔV = 2.25 km/s Orbit Design Source: Delta II Payload Planners Guide December 2006
  • 15. Three Stage Design Pros ◇ Reasonable ISP will meet our needs ◇ Better finert values Cons ◇ More complicated due to more staging ◇ Larger inert mass
  • 18. Choosing H202 and HTPB allowed us to have a reasonable ISP and good finert that will not be technologically challenging F2/H2 and O2/H2 had a large ISP, but too toxic. O2/RP-1- Needed oxygen tanks Propellant Trade Offs
  • 19. An initial mass of about 5500 kg, ISP of 285s and finert of 0.28 was chosen. The ISP of 285 allowed us to account for pressure losses in our system with the chosen propellant. Preliminary Sizing: Stage 1
  • 20. Helium pressurizes the propellant tanks to force the fuel and oxidizer tanks to the combustion chamber Pros: ◇ Simple, due to less components ◇ Easy to maintain Cons: ◇ Extra weight added because of the pressurant tanks Pressure-fed engine
  • 21. Injector in a liquid rocket engine mixes the fuel with the oxidizer to produce efficient and stable combustion This figure shows an injector designed with propellant valves (remote control) Injector
  • 24. Our design conditions were best met by ablative cooling The ablative material absorbs the heat as it ablates Cooling Type
  • 25. Pros: ◇ Simple ◇ Capable of stopping and restarting the engine, as long there is ablative material left Cons: ◇ Increase of weight ◇ Limited life in the engine (usually less than 2000 seconds) Typical ablative materials: Silica, Quartz, or Carbon Cloth and resin composites Cooling Type Trade offs
  • 26. Material Density (kg/m3) Ultimate Tensile Strength (GPa) Specific Ultimate Tensile Strength (Gpa/(kg/m3)) 2219 Aluminum 2800 0.413 15.04 Titanium 4460 1.23 28.81 4130 Steel 7830 0.892 11.23 Graphite 1550 0.895 58.88 SOURCE: Space Propulsion Analysis and Design, Humble, Henry, and Larson Pressure Tank Material
  • 27. Titanium has the best mechanical properties, but is difficult to work with and extremely expensive. Steel is cheap and easy to work with, but does not have the properties required for our design. Composites meet the property requirements and are lightweight. Our rocket cost increases, within reason. Material Trade offs Due to the addition of the pressurant tanks, composite materials were chosen to minimize the mass and therefore increase the ΔV This increased the cost of our rocket.
  • 29. ISP (s) (from RPA code) Initial Mass (kg) Propellant Mass (kg) Thrust (kN) Burn time (s) ΔV (m/s) (Actual) Stage 1 287 5550 3904 2698 4 3348 Stage 2 287 900 637 53 33 3317 Stage 3 287 200 109 12 26 2798 Total 9463 The actual ΔV, considering all the actual masses, is 2.5% bigger than design. Final Predicted Results
  • 34. Justifications 1. Small propulsion system to achieve an orbit 2. Uses current technologies proven to work 3. Performs a meaningful and desirable task

Editor's Notes

  1. Austin
  2. Austin
  3. Austin
  4. Austin
  5. Austin
  6. Austin
  7. Austin
  8. Cordarryl
  9. Cordarryl
  10. Cordarryl
  11. Austin
  12. Chris
  13. Chris
  14. Chris
  15. Chris
  16. Austin
  17. Uriel
  18. Uriel
  19. Uriel
  20. Uriel
  21. Cordarryl
  22. Cordarryl
  23. Lindani
  24. Uriel
  25. Joseph
  26. Joseph
  27. Joseph
  28. Lindani
  29. Lindani
  30. Lindani
  31. Austin