1 AIAA Space 2013
Lunar Lander Designs for Crewed Surface Sortie Missions
in a Cost-Constrained Environment
12 September 2013 | San Diego, CA | Revision A
Mark Schaffer Senior Aerospace Engineer, Advanced Concepts Group
[email protected] | +1.770.379.8013
2 AIAA Space 2013
Compare the performance, cost, and reliability of lunar
lander vehicle designs based on different propellant
combinations and vehicle configurations.
1. Introduction
2. Trade Study Definition
3. Assumptions
4. Results
5. Conclusions
Objective
Contents
4 AIAA Space 2013
Introduction to SpaceWorks
SpaceWorks is a responsive and nimble multidisciplinary aerospace engineering team
focused on independent concept design and systems analysis at fidelity levels suitable for
concept initiation through PDR.
We have over a decade of experience supporting advanced design and long range
planning activities for customers in private industry, NASA, DoD, and entrepreneurial
space organizations.
SpaceWorks is a privately held S-corporation. We are a small-business concern, are a
GSA PES schedule holder, and have a DCAA-approved accounting system.
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Rationale for Cislunar Space
SpaceWorks believes that cislunar space, i.e. the region of space surrounding the Earth and the Moon, is the next
logical step for NASA’s human space exploration program, with benefits in three areas:
Commerce – Development of a cislunar infrastructure will ensure continued U.S. leadership in the international
community, allow the U.S. to extend its economic influence beyond LEO, and enable the utilization of the
Moon’s material and energy resources.
Exploration – Cislunar space and the lunar surface provide a nearby proving grounds for new exploration
technologies and hardware; cislunar space is also a natural basing point for deep space missions.
Science – The study of the Moon’s surface and interior will be useful to the fields of planetary science and
solar system formation, and the lunar far side is of great interest to the astronomy community.
*Average distances based on mean Earth-Moon positions
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Propellant Options
Option Propellants Advantages Disadvantages
A LOX/LH2
• High performance
• Commonality with SLS upper
stage
• Hydrogen boil-off
• Low fuel density
B LOX/LCH4
• Low boil-off fuel and oxidizer
• Good fuel density
• Few heritage engines
• Low performance (compared
to LOX/LH2)
C LOX/RP
• Storable fuel,
• Low boil-off oxidizer
• Great fuel density
• Low performance (compared
to LOX/LH2 or LOX/LCH4)
D NTO/MMH
• Storable propellants
• Great propellant densities
• Lowest performance of all
options
• Toxicity necessitates special
handling
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Configuration Options
Option Configurations Advantages Disadvantages
1 Ascent + Descent
Stage
• Single shared propulsion
system
• Descent stage not reusable
2 In-Space + Lander
Stage
• Fully reusable
• Common propulsive elements
• LLO rendezvous
• Orbit plane change DV
3 Single Stage
Lander
• Fully reusable
• No LLO rendezvous
maneuver
• Single vehicle design
• Highest launch mass
• Largest physical lander size
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Design Assumptions
Habitat Element
2 crew for 20 days
20 m3 pressurized volume
Interfaces
• Two NASA suitlocks with spacesuits
• NASA docking system
Electrical power generated from ASRG
Open loop life support system
• Waste CO2 collected with LiOH canisters
• Waste water collected and tanked
Propulsive Elements
Existing or near-term technologies for structures, propulsion,
and subsystems
Passive thermal control for cryogenic propellants
Power provided by Habitat
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Liquid Rocket Engines
Parameter Option A Option B Option C Option D
Propellants LOX/LH2 LOX/LCH4 LOX/RP NTO/MMH
Engine Name CECE CECE RD-58M Aestus
Manufacturer Aerojet Rocketdyne Aerojet Rocketdyne RSC Energia EADS Astrium
Number per Stage 1 1 1 3
Cycle Expander Expander Gas Generator Pressure Fed
Vacuum Thrust 66.7 kN (15.0 klbf) 66.7 kN (15.0 klbf) 83.4 kN (18.5 klbf) 29.6 kN (6.6 klbf)
Vacuum Isp 460 sec 360 sec 349 sec 324 sec
Mixture Ratio (O/F) 5.8 3.6 2.5 1.9
Mass 210 kg 210 kg 300 kg 111 kg
Design / Modify Cost $260M $310M $77M $77M
Acquisition Cost $30M $31M $9.4M $16M
Mean Failure Rate 0.575% 0.575% 0.379% 0.097%
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Earth-Moon L1/L2
Configuration 1: Ascent + Descent Stage
3. L1/L2 to Lunar Surface
Transit Time = 3 days
6. Lunar Ascent
Ascent Stage
ΔV = 2,540 m/s
1. L1/L2 Loiter
Wait Time = 90 days
(before crew arrives)
Descent Stage
is Discarded
Moon
5. Lunar Surface
Stay Time = 14 days
2. L1/L2 Departure
Descent Stage
ΔV = 240 m/s
4. Lunar Descent
Descent Stage
ΔV = 2,790 m/s
7. Lunar Surface to L1/L2
Transit Time = 3 days
8. L1/L2 Arrival
Ascent Stage
ΔV = 240 m/s
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In-Space Stage
remains in LLO
LLO
Configuration 2: In-Space + Lander Stage
Earth-Moon L1/L2
3. L1/L2 to LLO
Transit Time = 3 days
9. Lunar Ascent
Lander Stage
ΔV = 1,900 m/s
1. L1/L2 Loiter
Wait Time = 90 days
(before crew arrives)
Moon
7. Lunar Surface
Stay Time = 14 days
2. L1/L2 Departure
In-Space Stage
ΔV = 240 m/s
6. Lunar Descent
Lander Stage
ΔV = 2,150 m/s
12. Lunar Surface to L1/L2
Transit Time = 3 days
13. L1/L2 Arrival
In-Space Stage
ΔV = 240 m/s
4. LLO Arrival
In-Space Stage
ΔV = 640 m/s
5. Separation 8. Plane Change
In-Space Propulsion
ΔV = 2,300 m/s
10. Rendezvous
11. LLO Departure
In-Space Stage
ΔV = 640 m/s
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Configuration 3: Single Stage
Earth-Moon L1/L2
3. L1/L2 to Lunar Surface
Transit Time = 3 days
6. Lunar Ascent
ΔV = 2,540 m/s
1. L1/L2 Loiter
Wait Time = 90 days
(before crew arrives)
Descent Stage
is Discarded
Moon
5. Lunar Surface
Stay Time = 14 days
2. L1/L2 Departure
ΔV = 240 m/s
4. Lunar Descent
ΔV = 2,790 m/s
7. Lunar Surface to L1/L2
Transit Time = 3 days
8. L1/L2 Arrival
ΔV = 240 m/s
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Vehicle Scale
1B 1C 1D
3A 3B 3C 3D
2A 2B 2C 2D
1A
Ascent + Descent Stage LOX/LH2 Ascent + Descent Stage LOX/RP Ascent + Descent Stage LOX/LCH4 Ascent + Descent Stage NTO/MMH
In-Space + Lander Stage LOX/LH2 In-Space + Lander Stage LOX/LCH4 In-Space + Lander Stage LOX/RP In-Space + Lander Stage NTO/MMH
Single Stage LOX/LH2 Single Stage LOX/LCH4 Single Stage LOX/RP Single Stage NTO/MMH
5.0m
Asc
ent
+ D
esce
nt
Sta
ge
In-S
pac
e +
Lan
der
Sta
ge
Sin
gle
Sta
ge
LOX/LH2 LOX/LCH4 LOX/RP NTO/MMH
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Vehicle Masses
0.0
10.0
20.0
30.0
40.0
50.0
60.0
70.0
Mas
s (t
)
Habitat Dry Stage 1 Dry Stage 2 Dry
Consumables* Stage 1 Propellant Stage 2 Propellant
* Includes crew, suits, crew consumables, vehicle fluids, and ACS propellant
Ascent + Descent Stage Lander + In-Space Stage Single Stage
1A 1B 1C 1D 2A 2B 2C 2D 3A 3B 3C 3D
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Non-recurring Cost
$0.0
$1.0
$2.0
$3.0
$4.0
$5.0
$6.0
$7.0
$8.0
$9.0
$10.0
Co
st (
$B F
Y13
)
Design, Development, Test, and Evaluation (DDT&E) Theoretical First Unit (TFU)
Ascent + Descent Stage Lander + In-Space Stage Single Stage
1A 1B 1C 1D 2A 2B 2C 2D 3A 3B 3C 3D
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Vehicle Reliability
0.0%
0.5%
1.0%
1.5%
2.0%
2.5%
3.0%
3.5%
4.0%
Co
ntr
ibu
tio
n t
o P
rob
abili
ty o
f L
oss
of
Mis
sio
n
Habitat Lunar Landing Engine Rendezvous Separation
Ascent + Descent Stage Lander + In-Space Stage Single Stage
1A 1B 1C 1D 2A 2B 2C 2D 3A 3B 3C 3D
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Comparison to Other Designs
Apollo ESAS Altair Option 1A
Number of Crew 2 4 4 2
Mission Duration 3 days 7 days 7 days 20 days
Number of Stages 2 2 2 2
Propellants NTO / UDMH LOX / LH2 LOX / LH2 LOX / LH2
Lander Mass 14.7 t 27.9 t 45.6 t 31.8 t
Vehicle Height 5.5 m 9.5 m 10.5 m 6.4 m
Airlock Height 3.0 m 5.5 m 7.0 m 4.0 m
Diameter 4.3 m 7.5 m 7.5 m 8.4 m
Maneuvers
(1) Descent from LLO
(2) Ascent to LLO
(1) Descent from LLO
(2) Ascent to LLO
(1) LOI
(2) Descent from LLO
(3) Ascent to LLO
(1) Descent from
L1/L2
(2) Ascent to L1/L2
10 m
0 m
5 m
ESAS and Altair Lander Images Credit NASA
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Results Summary and Ranks
Configuration Metric LOX/LH2 (A) LOX/CH4 (B) LOX/RP (C) NTO/MMH (D)
Ascent +
Descent Stage
(1)
Mass
(Initial in L1/L2)
31.8 t
2
36.7 t
3
39.5 t
5
47.4 t
9
Cost (DDT&E+TFU) $8,030
10
$7,520
6
$6,720
1
$7,400
4
Reliability
(LOM)
2.61%
8
2.58%
7
2.36%
4
2.18%
2
In-Space +
Lander Stage
(2)
Mass
(Initial in L1/L2)
29.0 t
1
37.4 t
4
42.0 t
7
55.3 t
11
Cost (DDT&E+TFU) $8,150
12
$7,530
7
$7,190
3
$7,670
9
Reliability
(LOM)
3.10%
12
3.08%
11
2.85%
10
2.66%
9
Single Stage
Lander
(3)
Mass
(Initial in L1/L2)
41.3 t
6
46.0 t
8
49.2 t
10
62.8 t
12
Cost (DDT&E+TFU) $8,150
11
$7,640
8
$6,760
2
$7,480
5
Reliability
(LOM)
2.43%
5
2.40%
6
2.18%
3
1.99%
1
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Observations
By taking advantage of the large fairing diameter available on SLS, the overall height of a lunar
lander can be reduced significantly compared to other designs
Compared to hydrogen, hydrocarbon or fully storable propellants can reduce vehicle size
significantly at the expense of increased mass
In-space + lander stage configurations are both less reliable and more expensive than the
ascent + descent stage configurations, but are fully reusable
Single stage options are likely infeasible due to high launch masses
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Conclusions
Option 1C, the RP two-stage ascent + descent vehicle, is the best overall performer
• The low cost and high reliability of the flight-proven RD-58M
• High density and zero boil-off properties of RP
Option 1B, the CH4 two-stage ascent + descent vehicle, is strong alternative to the RP vehicle
• Methane is synergistic with Mars exploration using ISRU
• LOX/LCH4 offers an appreciable performance advantage over LOX/RP
Differences in cost and reliability between RP and CH4 vehicles driven solely by engine: flight-
proven RD-58M engine vs. prototype CECE engine
Recommendation: Further development of exploration-focused CH4 engines would make
these options competitive in cost and reliability and benefit other exploration missions
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SPACEWORKS ENTERPRISES, INC. (SEI) | www.sei.aero | [email protected]
1040 Crown Pointe Parkway, Suite 950 | Atlanta, GA 30338 USA | +1-770-379-8000
27 AIAA Space 2013
Mass Details
Element Item LOX/LH2 (A) LOX/LCH4 (B) LOX/RP (C) NTO/MMH (D)
Habitat (Common) Inert 3.0 t 3.0 t 3.0 t 3.0 t
Crew/Consumables 1.0 t 1.0 t 1.0 t 1.0 t
Configuration 1: Ascent + Descent Stage
Ascent Stage Inert 2.6 t 2.2 t 2.3 t 2.6 t
Usable Propellant 6.3 t 7.5 t 8.0 t 9.4 t
Descent Stage Inert 2.6 t 1.8 t 1.8 t 2.1 t
Usable Propellant 16.3 t 21.2 t 23.3 t 29.3 t
Total System 31.8 t 36.7 t 39.5 t 47.4 t
Configuration 2: In-Space + Lander Stage
In-Space Stage Inert 2.0 t 1.9 t 2.0 t 2.5 t
Usable Propellant 10.1 t 14.6 t 16.6 t 23.2 t
Lander Stage Inert 2.6 t 2.6 t 2.9 t 4.2 t
Usable Propellant 10.2 t 14.3 t 16.5 t 21.4 t
Total System 29.0 t 37.4 t 42.0 t 55.3 t
Configuration 3: Single Stage
Lander Inert 6.7 t 4.8 t 4.9 t 6.0 t
Usable Propellant 30.7 t 37.2 t 40.3 t 52.8 t
Total System 41.3 t 46.0 t 49.2 t 62.8 t
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Cost Details
Configuration Cost LOX/LH2 (A) LOX/CH4 (B) LOX/RP (C) NTO/MMH (D)
Ascent + Descent
Stage (1)
DDT&E $7,390 $6,930 $6,210 $6,770
TFU $640 $590 $510 $630
Total $8,030 $7,520 $6,720 $7,400
In-Space +
Lander Stage (2)
DDT&E $7,430 $6,840 $6,560 $6,910
TFU $720 $690 $630 $760
Total $8,150 $7,530 $7,190 $7,670
Single Stage
Lander
(3)
DDT&E $7,460 $7,030 $6,190 $6,820
TFU $690 $610 $570 $660
Total $8,150 $7,640 $6,760 $7,480
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Reliability Details
Item LOX/LH2 (A) LOX/LCH4 (B) LOX/RP (C) NTO/MMH (D)
Configuration 1: Ascent + Descent Stage
Propulsive Maneuvers 0.68% 0.65% 0.43% 0.24%
Habitat 0.63% 0.63% 0.63% 0.63%
Lunar Landing 0.66% 0.66% 0.67% 0.67%
Rendezvous 0.45% 0.45% 0.45% 0.45%
Separations 0.18% 0.18% 0.18% 0.18%
Total LOM 2.61% 2.58% 2.36% 2.18%
Configuration 2: In-Space + Lander Stage
Propulsive Maneuvers 0.67% 0.65% 0.43% 0.24%
Habitat 0.63% 0.63% 0.63% 0.63%
Lunar Landing 0.66% 0.66% 0.66% 0.66%
Rendezvous 0.90% 0.90% 0.90% 0.91%
Separations 0.23% 0.23% 0.22% 0.22%
Total LOM 3.10% 3.08% 2.85% 2.66%
Configuration 3: Single Stage
Propulsive Maneuvers 0.68% 0.65% 0.43% 0.24%
Habitat 0.63% 0.63% 0.63% 0.63%
Lunar Landing 0.66% 0.66% 0.67% 0.67%
Rendezvous 0.45% 0.45% 0.45% 0.46%
Separations 0.00% 0.00% 0.00% 0.00%
Total LOM 2.43% 2.40% 2.18% 1.99%