A conventional Mercury sample return mission requires significant launch mass due to the large deltav required for the outbound and return trips, and the large mass of a planetary lander and ascent vehicle. Solar sailing can be used to reduce lander mass allocation by delivering the lander to a low, thermally safe orbit close to the terminator. Propellant mass is not an issue for solar sails so a sample can be returned relatively easily, without resorting to lengthy, multiple gravity assists. The initial Mercury sample return studies reported here were conducted under ESA contract ESTEC/16534/02/NL/NR, PI Colin McInnes, Technical Officer Peter Falkner. Updated solar sail capabilities were developed under the Ground System Demonstration program, funded by the NASA's In-Space Propulsion Technology (ISPT) Program

Adams, Charles L.

Montgomery, Edward E.

Young, Roy M.

NASA Technical Reports Server

Mercury Sample Return using Solar Sails
Montgomery1, Edward E.; Young1, Roy M.; and Adams2, Charles L.
1NASA Marshall Space Flight Center, Huntsville, AL; 2Gray Research Inc., Huntsville, AL
A conventional Mercury sample return mission requires significant launch mass due to the large Δv required for the outbound and return trips, and the large mass of a planetary lander and ascent vehicle. 
Solar sailing can be used to reduce lander mass allocation by delivering the lander to a low, thermally safe orbit close to the terminator. Propellant mass is not an issue for solar sails so a sample can be 
returned relatively easily, without resorting to lengthy, multiple gravity assists. The initial Mercury sample return studies reported here were conducted under ESA contract ESTEC/16534/02/NL/NR, PI 
Colin McInnes, Technical Officer Peter Falkner3,4,5.  Updated solar sail capabilities were developed under the Ground System Demonstration program, funded by the NASA’s In-Space Propulsion 
Technology (ISPT) Program.
1. Scoon, G., Lebreton, J-P., Coradini, M., et al, “Mercury Sample Return”, Assessment Study Report, ESA Publication SCI (99)1, 1999.
2. Oberto, B., et al, “Mercury Sample Return 2002-01,” Team X Advanced Projects Design Team Report, Jet Propulsion Laboratory, Pasadena, California, 2002.
3. McInnes, C. R., Hughes, G. W., and Macdonald, M., “Technical Note 3 – Mercury Sample Return,” ESTEC 16534/02/NL/NR, ESA/ESTEC Contract Report, University of Glasgow, 2003.
4. Hughes, G. W., Macdonald, M., McInnes, C. R., Atzei, A, Falkner, P., Terrestrial Planet Sample Return Missions Using Solar Sail Propulsion,” 5th IAA International Conference on Low-Cost Planetary 
Missions, ESA/ESTEC, The Netherlands, Sept. 24-26, 2003.
5. Hughes, G. W., and McInnes, C. R., “Mercury Sample Return Missions Using Solar Sail Propulsion,” International Astronautical Congress, Paper
IAC-02-W-2.08, Oct. 2002.
6. Murphy, D.M., and Murphey, T.W., “Scalable Solar Sail Subsystem Design Considerations,” AIAA 2002-1703, 43rd Structures, Structural Dynamics and Materials Conference, Denver, Colorado, Apr. 
22-25, 2002.
7. Murphy, D. M., “Validation of A Scalable Solar Sailcraft”, 53rd JANNAF Propulsion Meeting, December 2005.
8. Lichodziejewski, D., et al., “Vacuum Deployment and Testing Of a 20m Solar Sail System”, AIAA 2006-1705, 47th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials 
Conference, May 2006.
INTRODUCTION
Comparison of Solar Sail, Chemical and Electric Propulsion 
Options
Parameter SOLAR SAIL CHEMICAL SEP
Launch Vehicle H2A 202-4S Ariane 5E ** Atlas V 551
Launch Mass (kg) 2353 6500 ** 5775
C3 (km 2 s -2) 0 11.6 ** 7.8
Mission Duration (yrs) 4.4 7.2 ** 6.9 
** Chemical/SEP,
NASA’s In-Space Propulsion Technology (ISPT) Program is investing in technologies that have the potential to revolutionize the 
robotic exploration of deep space. For robotic exploration and science missions, increased efficiencies of future propulsion 
systems are critical to reduce overall life-cycle costs and, in some cases, enable missions previously considered impossible. The 
ISPT Program is developing technologies from a Technology Readiness Level (TRL) of 3 through TRL 6.
Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like sail 
made of a lightweight, reflective material. Because the Sun supplies the necessary propulsive energy, solar sails require no 
onboard propellant, thus reducing payload mass. With photonic pressure providing continuous thrust, sailcraft can conduct 
missions not available with conventional or electric propulsion:
• high-inclination orbital maneuver plane changes 
• fast flyby missions to the outer planets or extra-solar system
• orbit in non-Keplarian orbits (above the pole of a planet) and 
• hover indefinitely at a point in space
Over a two and one-half year period dating from 2003 through 2005, ISPT matured solar sail technology from laboratory 
components to full systems, demonstrated in as relevant a space environment as could feasibly be simulated on the ground.  Two 
20 meter square solar sail designs were manufactured, subjected to launch vibration and ascent vent profiles and deployed 
under high vacuum at NASA Glenn Research Centers’ Plum Brook Space Power Facility7,8. 
TECHNOLOGY DEVELOPMENT
SCIENCE OBJECTIVES
It is important to ascertain the surface age of Mercury to understand its geologic history, however accurate rock dating of Mercury surface samples is only possible on Earth. A high-
latitude landing site is required due to thermal constraints, and prior imaging of this site from the orbiter at a resolution of better than 1 meter per pixel is necessary. Even at high 
latitudes, landing in direct sunlight, or in permanent shadow would be undesirable. A landing site within a suitable crater, in partial shade, but with some light reflected from the crater 
walls is preferable, with a sample drilled from a rock outcrop within the crater.2 Baseline science objectives therefore are to acquire a surface sample of  350 g of surface regolith though 
a precision landing at a carefully selected high latitude landing site in partial shadow, within a suitably aged crater, with high resolution imaging for documentation during terminal 
descent. Sample pre-selection and pre-analysis will be conducted in-situ during landing site characterization using a robotic arm and small mobility device (20 m range).1
MISSION DESIGN
A baseline mission concept has been proposed which uses a 275 x 275 meter solar sail with an assembly loading of 5.9 g/m2 to deliver a lander, Sail Cruise Stage 
(SCS) and on-orbit science payload to a forced 100 x 7500 km Sun synchronous orbit at Mercury using a 2.85 year minimum-time trajectory, with a sail 
characteristic acceleration of 0.25 mm/sec2. From this initial orbit, the on-orbit science payload is used to survey potential landing sites for a period of up to 130 
days. The solar sail then descends to a 100 km polar orbit where a 1455 kg Mercury Descent Vehicle (MDV) is separated and descends to the south polar Bach 
region of the planetary surface using a blowdown bi-propellant system. A surface science phase of 4 days duration then takes place with the acquisition of 4 core 
samples and limited exploration using a small surface mobility device. Following sample acquisition, the Mercury Ascent Vehicle (again using a blow-down bi-
propellant system) delivers a small 15 kg Sail Rendezvous Vehicle (SRV) (cold gas system) containing the samples into a 100 km polar orbit. A dedicated 
rendezvous stage has been sized to avoid issues associated with the use of hot gas propulsion in close proximity to the sail film. After a 10 day hibernation in the 
100 km polar orbit to allow the solar sail orbit plane to match that of the rendezvous vehicle, final proximity maneuvering takes place along with sample 
transfer. The solar sail then spirals to escape, executes a 1.01 year minimum-time return trajectory to Earth and finally delivers the samples to Earth using 
aerocapture with a 17 kg ballistic Earth Return Vehicle (ERV) at parabolic approach speed.
Table 1: Sail Rendezvous Vehicle (SRV) system
sheet mass breakdown
14.5Total SRV Mass (Wet) 
1.1151.0Propellant for rendezvous 
13.4Total SRV Dry Mass 
0.11-System contingency 
13.312.2SRV Dry Mass 
0.43150.4Propulsion Mass (Dry) 
0.1150.1Propellant tank 
0.1150.1Valves, pipes 
0.23150.2Thrusters 
10.9109.8SRV Bus Mass 
2.2102.0Structure 
1.1101.0Thermal 
1.2101.1Telecomm
0.1100.1Mechanisms 
2.2102.0Power 
0.6100.5Command & data 
3.4103.1Attitude control 
2.0-2.0SRV Payload Mass 
2.0-2.0Sample container 
Total 
mass 
(kg)
Conting
ency 
(%)
Mass (kg)SRV Component 
Table 3: Mercury Descent Vehicle (MDV) system 
sheet mass breakdown 
1454.6 Total MDV Mass (Wet) 
960.0 15 834.8 Total Propellant Mass 
955.4 15 830.8 Propellant for Dv2 
4.6 15 4.0 Propellant for Dv1 
494.6 Total MDV Dry Mass 
4.9 1 -System contingency 
489.7 454.9 MDV Dry Mass 
162.5 15 141.3 Propulsion Mass (Dry) 
95.5 15 83.0 Propellant Tanks 
9.5 15 8.3 Valves, pipes 
57.5 15 50.0 Thrusters (5 of 6kN) 
149.4 10 135.8 MDV Bus Mass 
91.3 10 83.0 Structure 
3.3 10 3.0 Thermal 
0.0 10 0.0 Telecomm
24.2 10 22.0 Mechanisms 
9.7 10 8.8 Power 
4.4 10 4.0 Command & data 
16.5 10 15.0 Attitude control 
177.8 177.8 MDV Payload Mass 
2.9 -2.9 Surface instruments 
174.9 -174.9 MAV 
Total 
mass 
(kg) 
Contingency 
(%) 
Mass 
(kg) MDV Component 
Table 2: Mercury Ascent Vehicle (MAV) system
sheet mass breakdown
174.9Total MAV Mass (Wet) 
109.61595.29Total Propellant Mass 
109.01594.8Propellant for Dv2 
0.6150.5Propellant for Dv1 
65.3Total MAV Dry mass 
0.651-System contingency 
64.658.9MAV Dry Mass 
31.527.4Propulsion Mass (Dry) 
10.9159.5Propellant tank 
3.3152.9Valves, pipes 
17.31515.0Thruster 
18.61017.0MAV Bus Mass 
5.7105.2Structure 
2.2102.0Thermal 
0.0100.0Telecomm
0.6100.5Mechanisms 
2.5102.3Power 
2.7102.5Command & data 
4.9104.5Attitude control 
14.5-14.5MAV Payload Mass 
14.5-14.5SRV 
Total 
mass 
(kg)
Contingency 
(%)
Mass 
(kg)MAV Component 
Table 4: Sail Cruise Stage (SCS) system sheet mass 
breakdown
1905.3Total Sail Payload Mass 
401.810365.3SCS Bus Mass 
71.91065.4Structure 
55.01050.0Thermal 
27.11024.6Telecomm
177.110161.0Mechanisms 
44.21040.2Power 
11.01010.0Command & data 
15.51014.1Attitude control 
1503.51502.7SCS Payload Mass 
17.3516.5ERV 
31.6-31.6Science payload
1454.6-1454.6Lander 
(SRV/MAV/MDV) 
Total 
mass 
(kg)
Contingency 
(%)
Mass 
(kg)SCS Component 
Earth Arrival
September 8, 2018
Mercury Departure
September 3, 2017
X (AU)
.5
1.0
0
-.5
-1.0
Y (AU)
.5
1.0
0
-.5
-1.0
Z (AU)
.1
0
-.1
Mercury - Earth Return Trajectory 
Earth Departure
April 19, 2014
Mercury Arrival
February 24, 2017
X (AU)
.5
1.0
0
-.5
-1.0
Y (AU)
.5
1.0
0
-.5
-1.0
Z (AU)
.1
0
-.1
Earth - Mercury Trajectory 
A square solar sail is envisaged, using tip vanes for attitude control, sized to provide adequate slew 
rates for the planet-centered mission phases. The spacecraft (sail payload) is mounted centrally, within 
the plane of the solar sail, so that both faces of the core structure are free to be used as attachment 
points for the lander, and Earth return capsule. An outbound trip time of 2-3 years is desirable to be 
competitive with solar electric propulsion (SEP) and chemical mission trip times. This is enabled by a 
characteristic acceleration of 0.25 mm/s2. The chosen sail conceptual design used in this paper is based 
on the ATK Scaleable Solar Sail Subsystem (S4), since it can be extrapolated to large sail dimensions.6 
This design is based on coilable booms, and the boom linear density as a function of length can be 
combined with NASA/LaRC/SRS 2 or 5 μm CP1 film to obtain the sail assembly loading as a function 
of sail side length. It is assumed that conventional coatings are used, with aluminum (85% reflectivity) 
on the frontside and chromium (64% emissivity) on the backside. Shown below is the necessary sail 
assembly loading as a function of sail side length, for delivery of a 1905 kg spacecraft to Mercury with 
a characteristic acceleration of 0.25 mm/s2. It can be seen that the intersection of the 2 μm CP1 ATK 
S4 sail design curve with the 0.25 mm/s2, 100 km orbit payload curve yields the sail design point, with 
an assembly loading of 5.9 g/m2 and sail dimensions of 275 x 275 m. Included are updated curves
based on the 20 meter GSD for the ATK and L’Garde.
SAIL DESIGN
Manipulator 
Arm
SRV Sample 
Canister
MAV Pressurant 
Tank
MAV Engine
MDV 
Propellant 
Tanks
Landing Legs
SRV Cold Gas Tank
MAV Propellant 
and Oxidizer Tanks
MDV 
Oxidizer 
Tanks
Small Mobility 
Device
MDV Engines
Mercury Lander Concept
Stowed 
Booms
Earth Entry 
Capsule
Cruise Stage
Sail 
Storage 
Box
Rendezvous 
Vehicle
Ascent 
Vehicle
Descent 
Vehicle
Mercury Sample Return Launch Configuration
0 50 100 150 200 250 300 350 400
Sail Size (m)
20
18
16
14
12
10
8
6
4
2
S
a
i
l
 
A
s
s
e
m
b
l
y
 
L
o
a
d
i
n
g
 
(
g
/
m
2
)
Proton-K 
Block DZenit-3 SL
H2A202
4S
Dnepr 
Varyag
Soyuz ST 
Fregat
ATK S4 design 
with 2 μm CP1 
film
ATK S4 design 
with 5 μm CP1 
film
ATK GSD 
updated S4 
design with 2 
μm CP1 film
L’Garde 
GSD 
design 
with 2 
μm CP1 
film
https://ntrs.nasa.gov/search.jsp?R=20070002800 2019-08-30T00:21:37+00:00Z


Mercury Sample Return using Solar Sails

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