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.
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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
ENAE 791: Launch and Entry Vehicle Design 3
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
ENAE 791: Launch and Entry Vehicle Design 4
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
ENAE 791: Launch and Entry Vehicle Design 6
LV Trades: Fuel Types & Staging
Words, words, tables, words… Add Cost totals, maybe a table or
something
0
500
1000
1500
2000
2500
3000
3500
4000
4500
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
ENAE 791: Launch and Entry Vehicle Design 9
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 = $128.3 Billion 2012$
ENAE 791: Launch and Entry Vehicle Design 10
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
ENAE 791: Launch and Entry Vehicle Design 11
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
ENAE 791: Launch and Entry Vehicle Design 12
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
ENAE 791: Launch and Entry Vehicle Design 13
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
ENAE 791: Launch and Entry Vehicle Design 14
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
ENAE 791: Launch and Entry Vehicle Design 15
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
ENAE 791: Launch and Entry Vehicle Design 16
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%
ENAE 791: Launch and Entry Vehicle Design 17
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%
ENAE 791: Launch and Entry Vehicle Design 18
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
ENAE 791: Launch and Entry Vehicle Design 19
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
ENAE 791: Launch and Entry Vehicle Design 20
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
ENAE 791: Launch and Entry Vehicle Design 21
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)
ENAE 791: Launch and Entry Vehicle Design 22
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
ENAE 791: Launch and Entry Vehicle Design 23
TPBVP Burn Time = 242.5 (cont.)
Total Velocity
VFinal = 7.612 km/s
(@ 500 km)
Downrange Distance
Max ~ 500 km (X-dir)
ENAE 791: Launch and Entry Vehicle Design 28
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
ENAE 791: Launch and Entry Vehicle Design 29
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°
ENAE 791: Launch and Entry Vehicle Design 30
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
ENAE 791: Launch and Entry Vehicle Design 31
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
ENAE 791: Launch and Entry Vehicle Design 32
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
ENAE 791: Launch and Entry Vehicle Design 33
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
ENAE 791: Launch and Entry Vehicle Design 34
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
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