Ryan Mayes Duarte Ho Jason Laing Bryan Giglio. Requirements Overall: Launch 10,000 mt of cargo (including crew vehicle) per year Work with a $5M fixed

Ryan Mayes

Duarte Ho

Jason Laing

Bryan Giglio

ENAE 791: Launch and Entry Vehicle Design 2

Requirements Overall:

Launch 10,000 mt of cargo (including crew vehicle) per year

Work with a $5M fixed cost for operations/flight Launch Vehicle:

Minimize total program transport costAchieve a 500 km circular orbit

Crew Entry Vehicle:Maximize operational flexibility (L/D)Direct re-entry from 75,000 km HEOCapable of landing on ground


Assumptions

Overall20 year program lifeAll costing estimates in 2012 dollars

Launch vehicle85% learning curve for vehicle costingFor initial design, 9.2 km/s to LEO

Crew vehicleVehicle mass of 10,000 kgQuoted mass includes EDL systems


LV: Costing Trade Study

Base/Expendable ΔV = 9,200 m/s Stage Safe life > 30 flights

+100 m/s ΔV per Stage Reusable upper stage: +300 m/s ΔV Resulting ΔV Maximums

2 Stage = 9,700 m/s3 Stage = 9,800 m/s

All Costing and MER Analysis Completed in MS Excel 2007


LV: Costing Calculator (MS Excel)


LV Trades: Fuel Types & Staging

Words, words, tables, words… Add Cost totals, maybe a table or

something

0

500

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0 100 200 300 400 500

$/kg

Payload Mass Per Flight [MT]

1:C/E

1:C/E,2:C/E

1:St/E,2:C/E

1:C/E,2:St/E

1:St/E,2:St/E

1:So/E,2:C/E

1:So/E,2:St/E

1:C/E,2:C/E,3:C/E

1:St/E,2:C/E,3:C/E

1:St/E,2:St/E,3:C/E

1:So/E,2:C/E,3:C/E

1:So/E,2:St/E,3:C/E

1:C/Bal(30),2:C/E

1:C/Bal(500)+100,2:C/E

1:St/Bal(30),2:C/E

1:St/Bal(500)+100,2:C/E

1:C/Bal(30),2:C/Bal(30)

1:C/Bal(500)+100,2:C/Bal(500)+400

1:St/Bal(30),2:C/Bal(30)

1:St/Bal(500)+100,2:C/Bal(500)+400

1:C/Bal(30),2:C/Bal(30),3:C/E

1:C/Bal(30),2:C/Bal(30),3:C/Bal(30)

1:St/Bal(30),2:C/Bal(30),3:C/E

1:St/Bal(30),2:St/Bal(30),3:C/E

1:So/Bal(30),2:St/Bal(30),3:C/E

1:So/Bal(30),2:C/Bal(30),3:C/E

1:C/Bal(500)+100,2:C/Bal(500)+100,3:C/Bal(500)+400

1:St/Bal(500)+100,2:C/Bal(500)+100,3:C/Bal(500)+400


LV Trades: Modularity Effects


LV Trades: Safe Life Effects


Launch Vehicle: Costing Conclusions 2 Stages, Both LH2/LOX, Ballistic & Re-usable Upper: TPS, Parachutes, Legs, +100m/s ΔV for VL Lower: Parachutes, Legs, +100m/s ΔV for VL Payload is 50,000 kg to reasonably minimize cost

~ 200 launches per year ~ 2 weeks between flights of the same vehicle ~ 4 flights per week

Trades suggest lower cost for payloads above 50MT, but the greater required thrust negates any benefits and/or requires SRBs (3rd Stage)

641.68 $/kg 2012$ Total Lifetime Mission = $128.3 Billion 2012$


Engine Selection Launch:

S1 = 9 x Space Shuttle main engines (SSME/RS-25)

S2 = 1 x J-2X Re-entry: S1: 20 x P&W

CECE, S2 = 1 x J-2X Number of engines on

each stage was chosen to launch the maximum payload per launch in to orbit and maintain a mass margin of ~30% en.wikipedia.org/wiki/Space_Shuttle_Main_Engine


Launch Vehicle: Final Design

Total ΔV = 9,700 m/s Max. Payload = 50,000 kg Diam. 1 (Stage 1) = 10.2 m Diam. 2 (Stage 2) = 6 m Length = 80 m

D2

D1

L


Launch Vehicle: ΔV – Stages Target ΔV = 9,700 m/s St 1: ΔV1 = VE ln(m0/mf,1) St 2: ΔV1 = VE ln(m2/mf,2)

ΔV1 = 5,256 m/s ΔV2 = 4,444 m/s

Final Design Choice for Stage 2 ΔV


Launch Vehicle: Overview 2 Stage Uses LOX/LH2

propellant systems Total ΔV = 9,700 m/s

Stage 1 = 4,444 m/sStage 2 = 5,256 m/s

Total ΔV includes:9,200 m/s to orbit 300 m/s for reusables200 m/s for

deceleration components

Stage 1

Stage 2

Payload

Propellant 2

Propellant 1

Engine 2

Engines 1


Launch Vehicle: Stage 1 Total Propellant = 1,031,884 kg

Fuel (LH2) / Oxidizer (LOX) ratio = 6

Number of Engines = 9 SSME, 20 P&W CECE

Inert Mass fraction δ = .0914

Payload Mass fraction λ = .1953

Isp = 363 sec (SL)

LH2

LOX


Launch Vehicle: Stage 2 Total Propellant = 197,193 kg

Fuel (LH2) / Oxidizer (LOX) ratio = 5.5

Number of Engines = 1 J-2X Inert Mass fraction

δ = .1251 Payload Mass fraction

λ = .177

Isp = 448 sec (Vac)LH2

LOX


Launch Vehicle: Inert Mass Stage 1Component Mass (kg)

LOX Tank 9464

LOX Tank Ins 97

LH2 Tank 18869

LH2 Tank Ins 670

Payload Fairing 5099

Intertank Fairing 14104

Aft Fairing 1737

Launch Engines 31734

Component Mass (kg)

Re-entry Engines 3180

TPS System 0

Thrust Structure 4338

Gimbals 228

Avionics 1675

Wiring 3234

Landing Gear 3966

Parachutes 3305

Initial Estimate (Stage 1) = 132,208 kgFinal Inert Mass (Stage 1) = 101,699 kg

Final Design Margin = 30%


Launch Vehicle: Inert Mass Stage 2Component Mass (kg)

LOX Tank 1785

LOX Tank Ins 25

LH2 Tank 3883

LH2 Tank Ins 235

Payload Fairing 1178

Intertank Fairing 4099

Aft Fairing 1584

J-2X Engine 2472

Component Mass (kg)

TPS System 7070

Thrust Structure 494

Gimbals 93

Avionics 929

Wiring 1399

Landing Gear 1060

Parachutes 884

Initial Estimate (Stage 2) = 35,349 kgFinal Inert Mass (Stage 2) = 27,191 kg

Final Design Margin = 30.0%


Launch Vehicle: Analysis

Initial thrust/weight = 1.2 Stage 2 thrust/weight = 0.7 Assume constant mass flow rate (m_dot)

based on number of engines and all thrusters at full throttle

Thrust / weight ratio is a function of time; increases as propellant is burned.

Assume: Gravity; no drag Analysis performed in MATLAB using

integrated equations of motion


Launch Vehicle: Ascent First Pass Initial pitch angle:

89° (from horizontal)

Total Down Range after entire burn:

21 km

Down range distance of 2 km from the launch pad is achieved after 123 seconds

Down Range vs. Time


Launch Vehicle: Ascent First Pass Tstage,1:

215.7 sec Tstage,2:

49.8 sec Total Burn:

265.5 sec Final Height = 500 km This solution is not

optimized because final velocity is not totally in the x-direction

Altitude vs. Time


Launch Vehicle: Ascent TPBVP

Matlab solver: Two Point Boundary Value Problem (function: bvp4c.m)

Initial conditions:x = y = Vx = Vy = 0 km

Final conditions:y = 500 kmVx = Orbital Vel. @500 km

TPBVP solver in MATLAB creates the optimal trajectory to satisfy boundary conditions

Output: Min. flight time (saves cost)


Launch Vehicle: Ascent TPBVP

Stage 1 thrust scaled down to achieve an appropriate burn time

New Optimal Burn Time = 242.5 sec Indicates that another iteration

required to optimize

Altitude vs. Time

Final Velocity is fully in the x-direction for this optimal solution to the trajectory

Velocity vs. Time


TPBVP Burn Time = 242.5 (cont.)

Total Velocity

VFinal = 7.612 km/s

(@ 500 km)

Downrange Distance

Max ~ 500 km (X-dir)


3


2


1



Crew Vehicle: Costing Assuming:

Refurbishment rate of 3%

Nonrecurring cost for reusable vehicles doubled over expendable

1 crew vehicle for the program

Expendable vehicles cheaper up to 21st flight Reusable vehicles more cost efficient after 21st


Crew Vehicle: Lift and Drag Wanted cross range of roughly 2,000 km to span

entire continental US Drove selection for L/D = 1.3 Corresponds to angle of attack of 37.57° Newtonian Flow Estimations

CD,Sphere = 1

CD,Cone = 2sin2(δ)

Nominally chose CD = 1.3 as a baseline Based on Newtonian estimations and Soyuz figures

Sphere-cone with half angle δ = 54°


Crew Vehicle: Ballistic Coefficient Using parachutes necessitates that vehicle be at M = 1 or

lower at 3,000 m

β = 2000 kg/m2, vehicle area of 3.846 m2, diameter of 2.21 m


Crew Vehicle: Nominal Entry Trajectory Beta = 2,000 kg/m3

L/D = 1.3 FPA = -2° Downrange Max 35 km Peak velocity:

1.296 m/s at 30.3 km

Peak deceleration: 5.4483 g’s at 11.9 km


Crew Vehicle: Entry Heating

Heating rate approximation at stagnation point

Leading edge radius rLE = 3.298 m Max heating rate = 18.09 W/cm2

Total heat load = 470.94 J/cm2


Crew Vehicle: TPS Mass Estimation

AblativeHeuristic a function of

total heat loadQ = 470.94 J/cm2

TPS mass○ 2.18% of vehicle mass

Reusable (Shuttle tiles)Small sample size, no heuristicsMass was scaled based on shuttleTPS mass

o 8.63% of vehicle mass


Crew Vehicle: Landing Drogues upon entering atmosphere to stabilize, parachutes

employed as final reentry phase M = 1 achieved at roughly 3000m as a result of β selection;

allows for parachute deployment Parachute radius of 10m: terminal velocity of roughly 10 m/s 3 Parachutes: Loss of 1 chute results in a 20% terminal

velocity increase.

0 2 4 6 8 10 12 140

50100150200250300350400450

Chute Radius (m)

Chute Radius (Three count) (m)

One Loss Velocity

Impact Velocity (m/s)

Rad

ius(

m)

V 2mg

Sd

S Aequivalent 3r2


Crew Vehicle: Landing


Crew Vehicle: Landing

Documents

Ryan Mayes Duarte Ho Jason Laing Bryan Giglio. Requirements Overall: Launch 10,000 mt of cargo (including crew vehicle) per year Work with a $5M fixed