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