Dankanich, John

Woodstock, Gordon

NASA Technical Reports Server

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Growing Technology Base for SEP
•
• Solar-electric propulsion (SEP) is becoming of interest for application to a wide
range of missions.
• Benefits of SEP are strongly influenced by system element performance, especially
the power system.
• Solar array performance is increasing rapidly and promises to continue to do so for
another 10 to 20 years (graph)
• Cost per watt is decreasing.
• Radiation hardness is increasing.
RESEARCH SOLAR CELL EFFICIENCIES & PREDICTIONS
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New concepts for designing
SEP are emerging. These
lead to changes in the best
ways to apply SEP
technology to missions, and
broadening of the practical
uses of SEP technology
compared to competing
technologies.
YE:'R
National Aeronautics and Space Administration 2
Why Use SEP?
•
• Application of SEP technology is favored when:
• (1) the mission is compatible with low-thrust propulsion: no need to
produce accelerations comparable to planetary gravity fields and
adequate time to achieve the mission ideal delta V;
• (2) the mission needs high total delta V such that chemical propulsion
is disadvantaged; and
• (3) performance enhancement is needed to make the mission
compatible with existing launch capabilities, or to provide or cost
reduction.
National Aeronautics and Space Administration 3
SEP Mission Regimes
•
Mission Regime Representative SEP Applications
Regime 1 - Human Earth Orbit, near-Earth Orbit adjust and makeup, including application to
missions (includes ISS) ISS
Regime 2 - Robotic LEO to near planets; Earth & Robotic LEO to higher Earth orbits; Mars sample
space observation, planetary science and return, missions to NEOs, Venus
sample return
Regime 3 - Human HEO to Lunar; space transfer Cargo mission support for human lunar and
humans/cargo with landing and bases other Earth vicinity missions.
Regime 4 - Robotic near Sun; includes Mercury Near-sun or out of ecliptic missions requiring
and solar probes/polar missions high delta V; trade vs solar sails
Regime 5 - Human inner planets; Mars ex- Humans and cargo to Mars and NEO and main
ploration & landing, asteroids, space transfer belt asteroids.
Regime 6 - Robotic outer planets; orbiters, All of these except landers and ascent vehicle
probes, landers and sample return missions themselves.
Regime 7 - Human outer planets; Jupiter and Perhaps some portions of trajectories but these
Saturn and their moons, landing and return missions need thrust far from Sun.
Regime 8 - Robotic beyond planetary system; Very high performance SEP possibly can inject
Kuiper Belt, Oort Cloud, interstellar missions these by producing enough delta V near Sun.
National Aeronautics and Space Administration 4
Effects of Elliptic Starting Orbits
•
4000010000 20000 30000
Apogee Alt
Rep. LlV Perf (Atlas 401)
-+- Calculated,
•
Perigee 500 km
,
_ Quote, Perigee
~ 170 km
\\ • Quote, Circular
\~ ~ Calculated,Circular
\ ~~I
~ -----. I4000
o
5000
9000
10000
0) 8000
..lIl::
'0
~ 7000
>.
co
a.. 6000
-+- Ending Inclination
_ Time Below 1200 km
--.- Time Below 10,000 km
~ Total Time
Starting Apogee Altitude km
a 5000 10000 15000 20000 25000 30000 35000
Elliptic Start LEO-GEO
'-'-1
0.1 +----+----+----t----+---+--+------{
1000
11l
*
>-
l'O 100
"0
Q)
E
I-
06 10
Ol
Q)
"0
s::::
U 1
s::::
• Low circular starting orbit causes lengthy exposure to van Allen belts and
significant periods in the LEO debris environment.
• Launch vehicles pay a high performance price for higher circular orbits
• Elliptic starting orbits payoff in reducing radiation and debris exposure time
• For all-chemical propulsion the GEO/LEO payload ratio is about 20%; even
with an elliptic starting orbit, SEP is expected to double that
National Aeronautics and Space Administration 5
Orbit Raising Thrusting Strategy
•
• Two questions: (1) thrust all the time or part of the time, and (2) thrust in the direction of flight or in some
other direction?
• Apogee "gap", period of non-thrusting centered on apogee; thrusting at apogee contributes least to orbit
raising. (Thrust pointed in direction of flight. About 15 days of added trip time reduced delta V by a few
hundred mls and could increase payload several percent.
• Irregularity of data caused by ellipticity of orbit and difference in contribution to apogee radius between
perigee and apogee thrusting.
Gap Study Delta V Trip Time Occultation Fraction
y =: -0.3494x2 + 51.193x + 1173.3
R2 =: 0.6943
302510 15 20
Gap, Degrees
5
1.1
1.2 -,----,-------,----,-----,-----.----,
-c::
a> 0.9(.)
..
a>
a-
0.8
0.7
115 0.6
0
110105100
Trip Tim e days
95
3100
3000
2900
~
E 2800
>
III 2700
-Qj
0
2600
2500
2400
90
National Aeronautics and Space Administration 6
Plane Change During Ascent
•
• Orbit raising is often from an inclined starting orbit to zero inclination GEO.
• One can optimize how plane change is effected. We seek a control law practical for spacecraft GN&C. We
find: tan J3 =Ap cos S/(2Aaa) where A)? is a constant and Aa varies with a (semi-major axis). S is angle around
the orbit from ascending node and p is yaw angle.
• Tried forms 1/(Aa a) = kau where k and a are adjusted to give minimum /).V and desired plane change, and Aa
= k1 exp (k2 aU), same general idea.
• Exponential power law is very slightly better than the simple power law
Delta V LEO-GEO LEO-GEO Guidance Laws
"-
-+-power law
_ exponential power~~ law-.-exp pwr law 0.5
')~~ -x-exp pwr law 0.75
~~~ /
~~~~ /
I
--.-.
sma/smaO exponent
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•
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• Power Law
• Exp Power Law
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ai'
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c: 10ro
0-
5
o
o 10000 20000 30000
Semi-Major Axis, km
40000 50000
National Aeronautics and Space Administration 7
Plane Change with Elliptic Starting Orbit
•
• For elliptic starting orbit, eccentricity is enough that we need a yaw steering law that includes eccentricity
and argument of periapsis.
• Derivation uses coordinate transformations. Best one to formulate the problem is orbit coordinate system
where x axis points through line of nodes instead of periapsis.
• Used a momentum vector approach, where instantaneous plane change is given by ratio of (1) orbital
momentum vector rotation around the line of nodes due to thrust, to (2) current orbital momentum vector.
The result is:
Orbit Parameters (assumed)
Yaw Steering
80 -
_Arg P = 0
70 - _A rg P = 30 -t---------,r-,iPIod-------t----------j
11594
6 0 +_----__+--~------"'~-_____R1n_-__+---_____h_I81._e
0.41262722
o +_-----=~______,~-----+_---a------"_ __+----____J
-10 ~~~=-__L------l-~~~~~~-~
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QI
...
g'40+-----_+--,r--+----~.______+----..rr__-_+---~
"0
~~ 30
e 20 +_-------tt.--F----+_--k---+---'-"'-'-'-'-"""-"'--¥---------...,
~
~ 10+-----a4-1-----+-------1.---'\,---_+-------I
New Terms:
CD argument of periapsis
l.l Earth geopotential
V SEP velocity
T1 1 + e cos 8
T2 sqrt(1 + 2e cos 8 + e2)
-20 +-----_+-----+-------+----____J
o 100 200 300 400
True Anomaly (Theta) degrees
National Aeronautics and Space Administration 8
Further Work Needed:
•
• When apogee reaches GEO, must hold it or otherwise circularize the orbit at GEO altitude.
• Apogee hold pitch modulation works but is inefficient. Blending the apogee hold (uses some fraction of
apogee hold pitch) is better.
• Should derive a guidance law where eccentricity is also represented in the Hamiltonian so that
circularization can be optimized.
• Semi-Major Axis
• Periapsis
~-Apoapsis
Blended Apogee Hold and Plane Change
45000 -,-------,-----------,-------,
150100
• Semi-Major Axis
• Perigee
-A-Apogee
50
40000
35000
"'
30000
...
Ql
.... 25000Ql
E
0 20000
S2
15000
10000 -
5000
a
015010050
Orbit Dimensions
45000
40000
35000
30000
25000
E
~ 20000
15000
10000
5000
0
0
Time days Time days
National Aeronautics and Space Administration 9
Isp Optimization
•
• For electric propulsion, with limits on operating time, there is an optimum Isp. Too low Isp,
too much propellant consumption. Too high, too much SEP inert mass.
• The graph illustrates with a "simple" escape trajectory not using a planet swingby. A
minimum TIW is required to avoid thrust from falling off faster than mass in a spiral away
from the Sun. Delta V assigned, 50 km/s, was to occur after Earth escape to produce solar
system escape with significant residual velocity.
• The chartillustrates optimum electric propulsion Isp. The required mass factor, 10 kg/kWe or
less for an entire SEP, is very difficult.
SEP Isp Optimization Example
- ---
.--
!"'_-.,1I
) 1I
~ .............lifT"''''
.~" ~ "Y ....~~
71
I ......- Payload = SEP
•
Inert
_ Payload 10% of
Start Mass
9
8
Q)
~ 7
.::&.
tii 6
.::&.
~ 5
.0
'iii 4
III
E 3
I-
Q) 2
D..
1
a
a 2000 4000 6000 8000 10000 12000
National Aeronautics and Space Administration
Specific Impulse (Isp)
10
Lunar Science Application
•
• Mission profile, useful for ISRU experiments, e.g. oxygen production.
• Combined SEP-chemical propulsion system delivered to low Earth orbit.
• The chemical system propels the vehicle to an elliptic Earth orbit, to reduce trip time and radiation/ debris exposure.
• SEP takes over and delivers the vehicle to a lunar intercept.
• SEP or chemical system does lunar capture to a low lunar orbit.
• The chemical system performs lunar landing.
• Upon landing, the SEP solar array is re-deployed to power the surface mission.
• Modest performance augmentation (not analyzed) obtained by jettisoning SEP thrusters and
power processors before landing.
• Tradeoff looked at how much delta V should be delivered by SEP.
• Chemical was assumed to use hydrogen-oxygen with lsp 450.
• Delta IV Heavy delivery to the starting elliptic orbit; payload to a typical starting orbit 250 x
10,000 km is about 16,800 kg.
• Delta IV could deliver to TLI, but we assumed lunar lander chemical propulsion provided all
chemical delta V beyond starting orbit.
• EP run time held constant 210 days by adjusting power level.
• We assumed power required on the surface is 25 kWe. If 25 kWe yielded trip time less than
210 days we accepted that.
National Aeronautics and Space Administration 11
Lunar Science Application Results
•
• SEP Isps 1500 and 2000. Minimum chemical delta V occurs with a 1500 m/s boost to elliptic orbit and 2100
m/s for lunar landing. Maximum occurs with a 3000 m/s boost to an elliptic orbit with apogee near lunar
distance, most of the .Iunar orbit insertion, and the landing. SEP delta V ranges from a minimum of 900 m/s
to a maximum of 6010 m/s.
• Maximum payload, minimum cost with maximum use of electric propulsion
• Three final caveats:
• (1) complex trade, present result is a trend analysis only.
• (2)Assumed equal value for any kg of payload. Great value only if a smaller launch vehicle can be used.
• Cost estimates were crude; result is sensitive to cost of electric propulsion hardware. If solar electric propulsion is more costly
relative to other costs there may be no cost benefit to this particular use of electric propulsion.
Electric Power Payload Trade Cost Trade
10
o
3000 4000 5000 6000
Chemical Delta V m/s
__ Isp 1500
---Isp 2000
-lr- All Chem ----+------'<-+---1
100000 +-----+------l-----l----.J
3000
150000 ,_---~---,_---,_____,
'0 140000
1'0
o
>-1'0
ll. 130000 +-----+-----+-----+--4----1
OJ
u
I:
OJ
~ 120000 -j-----Yrif---+--~~----t-''''_I
til
.:.:.
......
iA- 110000 +_-~-+_---+_---+------I
4000 5000 6000
Chemical Delta V m/s
2000 +_---+----+-----t-----.J
3000
2500 +-----+----+----+-----I
4500 ,-----,-------,------,------,
4000 +---~+-----+--
~ 3500 +-------f'\:-----"-~+_---+_"III---j
"C
1'0
o
>-&. 3000 -!----+----+--~--t-----l
6004000 5000
Chemical Delta V m/s
__ Isp 1500
...........
---Isp 2000
~
~ ~
~"...
" '"
......... ,
20
80
70
90
u
E 60
u
OJ
W 50
~ 40
~
:: 30
:.::
National Aeronautics and Space Administration 12
500 kWe SEP Concepts
Conventional
Modular, Distributed Thrusters
National Aeronautics and Space Administration
•
13
Lunar Missions Requiring Heavy Cargo
•
• SEP can deliver an LSAM-derived cargo lander to lunar orbit and return it to Earth after it
unloads the cargo on the surface and returns to lunar orbit. LSAM would use pump-fed LOX-
methane or LOX-LH2 for descent and ascent.
• Four options analyzed:
• (1) Conventional expendable one-way chemical propulsion delivery.
• (2) Replaces translunar stage with 500 kWe SEP; SEP and the cargo lander are now re-used; cargo lander,
returns empty to lunar orbit.
- (3) Supply a modest amount of
lunar oxygen to help return the lander
to lunar orbit.
- (4) Provide (after the first landing)
all of the lander oxygen (after the first
landing). Load oxygen on the lunar
surface, fuel from Earth in lunar orbit.
- All cases land 10 - 11 t. of cargo.
Launch requirements incrementally
reduced with each technology
advance.
National Aeronautics and Space Administration
90000
80000
Cl
~" 70000
If)
~ 60000
"5 50000
c:
i6 40000
...J
-g 30000
...
g. 20000
Ql
a:: 10000
o
~
-
~
- I-- f-------
- I-- I-- I-- -
- I--- f------- I-- -
- I--- f------- I--- -
Chern LEO SEP LEO SEP LEO SEP LEO
delivery Deli\€ry Deli\€ry Deli\€ry
LLOX LLOX (2)
14
Planetary Science Missions
•
• Launch vehicle high-thrust propulsion has high leverage deep in Earth's gravity well;
optimization increases C3 until SEP leverage due to its high Isp exceeds the high-thrust
leverage. Graph shows the relative leverage of high thrust and SEP propulsion versus C3.
If one starts optimization at C3 0 or greater, the result will be as noted, C3 10- 15. However,
there is a region at C3<O where the SEP also has good leverage.
_La Thrust
--fr- La Thrust
Escaped
• A simple test calculation showed that a
SEP starting from elliptic orbit with C3 --
-35 and spiraling to escape could
deliver more payload than using the
same launch vehicle to launch a smaller
SEP to C3 -- 15. The cost trade
depends on SEP cost delta for larger &
more power versus launch cost.
C3 Sensitivity
40 -
____ Hi Thrust
>
I'll
....Qj
o
~
I:
:: 20 +-----~";i--------------1
QI
D.
.!:
I'll
I'
M
U
10050a
C3
-50
a +-------,----~~-___,____----1
-100
National Aeronautics and Space Administration 15
Planetary Exploration
• SEP and other new technologies have high leverage for reducing human Mars mission
transportation cost. Best mission profile we have found:
1. SEP delivers ITV* to
elliptic Earth orbit,
fueled with hydrogen;
resupplies hydrogen for
subsequent missions
2. SEP delivers LOX
tanker to
lunar orbit
7. Crew DE&L
1
National Aeronautics and Space Administration
3. LOX tanker shuttles
between lunar surface and
elliptic orbit.
4.
5.
6.
2,2.1
2.1 SEP delivers LH2 to
lunar orbit to replenish
LOX tanker
lTV goes to elliptic Mars orbit,
chemical trans-Mars injection.
Aerocapture into elliptic Mars
orbit
lTV returns to elliptic Earth
orbit, chem TEI**, aero-
capture EOI***.
*1 nterplanetary transport vehicle
**Trans-Earth injection
***Earth orbit insertion
16
Mars Mission Cost Results, Per Trip
•
• Elliptic Earth parking orbit enables interplanetary vehicle to return with its aerobrake and
transfer habitat for re-use (2nd bar).
• Diagram of previous chart adds lunar oxygen, middle bar below.
• Production of LOX-methane propellant on Mars enables re-usable Mars lander/ascent
vehicle (MLAV) based on Mars. Ascent of entire vehicle on Mars propellant; descent using
propellant from Earth.
• Low-cost launch represents a partially re-usable launcher, flyback booster with expendable
LOX-LH2 upper stage.
Add Add low-
reusable cost launch
MLAV
EEOjSEP EEOjSEP
plus lunar
LOX
Baseline
Cryo-
Aerocap
Cost Comparison
12000 -,-------------------,
~ 10000 -
~
c:
o 8000
'iii
Ul
~ 6000-
I.
~ 4000
...
Ul8 2000-
O-l---''------''--''----r------'---'-~"'"'--''----..____'----..J'___r_----'-----''--,_____'_-----''--l
SEP for delivery from LEO
to elliptic orbit is 500 kWe
- 1 megawatt. More cost
effective than a multi-
megawatt SEP going to
Mars and back because
non-recurring investment
is much less.
•
National Aeronautics and Space Administration 17
Acknowledgement
The work described in this paper was funded by the In-Space Propulsion
Technology Program, managed by NASA's Science Mission Directorate in
Washington, D.C., and implemented by the In-Space Propulsion Technology
Project at Marshall Space Flight Center in Huntsville, Alabama. The program
objective is to develop in-space propulsion technologies that can enable or
benefit near and mid-term NASA space science missions by significantly
reducing cost, mass or travel times.
National Aeronautics and Space Administration
•
18
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Application of Solar-Electric Propulsion to Robotic and Human Missions in Near-Earth Space

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