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Rapid Response Launcher System (RRLS)
Austin
Ramirez, UrielJohnson, LindaniGonzalez, Austin Solomon-Williams, Cordarryl M. Smith, JosephKnill, Christian
Austin
Problem StatementA rocket system to set a network of Search & Rescue Satellites
?
Austin
““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.”
Austin
Payload - 100lb S&R Satellite
Orbit - Circular Sun Synchronous at an altitude of 165 km
Launch Location - Poker Flat Research Range, AK65.1167° N Lat, 147.4612° W Lon, 436.7 m above sea level
Mission RequirementsAustin
Extraneous RequirementsEnvironmentally Friendly Propellant System
Total system weight less than 10,000kg
Minimum orbital lifetime 6 months
Austin
Missing PlanesAustin
Poker Flat Research Range, AK
MH 370
Orbital Elements from STKPeriod 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
cordarryl
Satellite Tool Kit Simulation Verification(2D Graphic of Earth)
cordarryl
Satellite Tool Kit Simulation Verification(3D Graphic of Earth)
cordarryl
Preliminary Design?
chris
ΔV Design 1Burnout Velocity -
Velocity of Launch site -
chris
ΔV Design 2Velocity Needed -
Design Velocity -
Assuming losses of 0.9km/s
chris
Three stages:Stage 1 will produce a ΔV = 3.5 km/sStage 2 will produce a ΔV = 3.5 km/sStage 3 will produce a ΔV = 2.25 km/s
Orbit Design
Source: Delta II Payload Planners Guide December 2006
chris
Three Stage DesignPros
◇ Reasonable ISP will meet our needs
◇ Better finert values
Cons◇ More complicated due
to more staging◇ Larger inert mass
chris
System Level Performance?
Propellant ChoiceUriel
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 OffsUriel
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 1Uriel
Helium pressurizes the propellant tanks to force the fuel and oxidizer tanks to the combustion chamberPros:◇ Simple, due to less
components◇ Easy to maintain
Cons:◇ Extra weight added
because of the pressurant tanks
Uriel
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)
cordarryl
Injector
Injector Types
AUTOCAD DRAWING Nozzle efficiency 97 %
Lindani
Nozzle Sizes
Our design conditions were best met by ablative coolingThe ablative material absorbs the heat as it ablates
Lindani
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 58.88SOURCE: Space Propulsion Analysis and Design, Humble, Henry, and Larson
Joseph
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 offsDue to the addition of the pressurant tanks, composite materials were chosen to minimize the mass and therefore increase the ΔVThis increased the cost of our rocket.
Final Predicted Design Results?
Joseph
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 9463The actual ΔV, considering all the actual masses, is 2.5% bigger
than design.
Joseph
Final Predicted Results
Altitude SimulationLindani
Velocity SimulationLindani
Mass SimulationLindani
ConclusionTheoretically, it is feasible. ?
Austin
Justifications1. Small propulsion system to achieve an orbit
2. Uses current technologies proven to work
3. Performs a meaningful and desirable task
THE END