Representative interplanetary space vehicle systems are sized to compare and show sensitivity of the initial mass required in low Earth orbit to one mission mode and mission opportunity. Data are presented to show the requirements for Earth-Mars opposition and conjunction class roundtrip flyby and stopover mission opportunities available during the time period from year 1997 to year 2045. The interplanetary space vehicle consists of a spacecraft and a space vehicle acceleration system. Propellant boil-off for the various mission phases is given for the Lox/LH (Liquid Oxygen/Liquid Hydrogen) propulsion systems. Mission abort information is presented for the 1999 Venus outbound swingby trajectory, transfer profile

Young, Archie C.

English

NASA Technical Reports Server

N87-17729
MISSIOBMIDV]_ICI_ SIZIN6 _ITIVITI_
Archte C. Young
Marshall Space Flight Center
Marshall Space Flight Center, AL 35812
ABSTRACT
Representative Interplanetary space vehicle systems are sized to
compare and show sensitivity of the initial mass required in low Earth
orbit to one mission mode and mission opportunity. Data are presented to
show the requirements for Earth-Mars opposition and conjunction class
roundtrlp flyby and stopover mission opportunities available during the
time period from year 1997 to year 2045. The interplanetary space
vehlcle consists of a spacecraft and a space vehicle acceleration system.
Propellant boil-off for the various mission phases Is given for the
Lox/LH (Liquid Oxygen/ Liquid Hydroge_ propulsion systems. Mission abort
Y
information is presented for the 1999 Venus outbound swingby trajectory.
transfer profile.
INTRODUCTION
This paper presents information on performance and operational
requirements and their sensitivity to flyby, Venus swlngby with stopover,
and conJuction class missions to Mars with stopover misslons. The time
period considered in developing this information is 1996 to 2045. The
initial mass required in low Earth orbit was determined for each launch
opportunity associated with the three classes of missions. The Mars
flyby is a nonstop encounter with Mars; the Venus swlngby mode
opposltion-class mission is a mission of less duration than the con-
Junction class mission but only allows a short stopover time of 60 days
at Mars. Conjunction class missions require longer stopover times, up to
550 days, at Mars, but require less propellant.
Information developed in this paper is not final, as configurations
of the transportation vehicles are not firm. Different values of the
Mission Module (MN), Mars Excursion Module (MEM), and Mars probes may
appear. The important thing to note is the relative comparative values
presented for the different mission modes.
ASSUMPTIONS
Pertinent assumptions used in this study are given for the departure
and capture orbit parameters, propulsion stages and planetary spacecraft
87
https://ntrs.nasa.gov/search.jsp?R=19870008296 2020-03-20T12:30:34+00:00Z
elements (Figure 1). The Interplanetary space vehicle was assumed to be
assembled in, and depart from the 270 nm altitude, 28.5 degrees Inclina-
tion, Space Station circular orbit. For the all propulsive case, re-
quired interplanetary velocity increments are achleved by three propul-
sive stages. The first propulsion stage effects the Earth escape
maneuver, the second stage brakes the spacecraft and Earth braking stage
into the Nars elliptical capture orbit and effects the escape maneuver
from the Nars elliptical orblt. The third propulsion stage brakes the NN
into a 24-hr elliptical orbit at Earth return. Each of the three pro-
pulslon stages' mass fractions were developed using scallng equations.
For the Nars aerocapture and Earth return aerobraked case, the
Interplanetary velocity Increments are achieved by two propulsive stages.
The first and second stages were used to effect the Earth and Nars escape
maneuvers, respectively.
Venus swlngby, outbound, inbound, or double swtngby, was used to
lower the energy required for the Nars opposition class missions. The
Venus closest approach distance was constrained to be equal to or greater
than 0.1 planet radii (330 nm).
For the conjunction class missions, type I (<180 deg) or type II
(>180 deg) Hohmann transfer trajectories were used. The Nars stopover
tlme was optimized to achieve minimum lnitlal weight in Earth orblt.
Interplanetary trajectory parameters (launch dates, trip tlmes,
heliocentric transfer angles, etc.) have been determined which result In
a mlnlsum total initial weight to be assembled in the Space Statlonts
orbtt. The varlable propulsion stages were sized uslng general scallng
weight laws which are dependent upon propellant loadlng. These coeffi-
cients are input to the Interplanetary trajectory shaping program. Up to
six major Interplanetary maneuvers can be optimized.
INTERPLANETARY SPACE VEHICLE
The spacecraft is made up of a NN (the livlng and work area for the
crew), a NEH and experimenter accommodations. A number of unmanned
probes and orblters are included to complement the manned activity.
NaJor elements of the spacecraft are interconnected by pressurized
tunnels allowing shirt sleeve passage between them. A minimum crew of 6
Is necessary to operate the space systems and perform a reasonable scien-
tific exploration program.
88
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Two Interplanetary space vehicle configurations for the opposition
class mission via an outbound Venus swlngby for the year 1999 opposition
opportunity are given in Figure 2. Information for each of the propul-
sion stages and the total interplanetary vehicle weight is given. The
total initial mass required in the Space Station orbit for the all pro-
pulsive configuration is 3,575,321 lb; for a configuration that utilized
aerobraktng at Mars capture and Earth return, the total initial mass
required in the Space Station orbit is 1,433,294 lb.
Earth return with aerobrake entry has been analyzed and results show
that with an Earth return C3 greater than 25 km2/sec the g-load wlll be
in excess of 5 g's. This high g-load probably cannot be tolerated by the
crew. Earth return with C 3 greater than 25 km2/sec 2 will require propul-
sive braking in order to stay within the g-load constraint.
MISSION AND VEHICLE SIZING SENSITIVITY
In mission profile design and vehicle slzing there are many
variables which influence the resultant mission profile and space vehicle
configuration. Some of the more significant variables include: (1)
Earth launch window duration, (2) Stay time at Nars, (3) _ weight, and
(4) MEM weight, including Mars lander capsule weight and Mars ascent
capsule weight.
Sensitivities to the Earth launch window duration and Mars slay time
for the 1997 and 1999 opportunity Venus swlngby mission profiles Is given
in Table 1. For the 1997 opportunity, a 40 day stay time at Mars and an
Earth launch window of 10 days requires 1,591,700 pounds initial welght
in low Earth orbit to perform the mission. A 60 day stay tlme and an
Earth launch window of 30 days wlll require 1,949,700 pounds of initial
weight in low Earth orbit to perform from the 1997 launch opportunity;
this weight ls an Increase of 22 1/2 percent over a 40 day stay time and
10 day launch window case. The 1999 launch opportunity ls not as
sensitive as the 1997 opportunity. A 60 day stay time and a launch of 30
days requires an initial weight in low Earth orbit of 1,434,200 pounds;
this weight Is an Increase of 6 3/4 percent over a 40 day stay time at
Mars and a 10 day Earth launch window.
The interplanetary space vehicle sensitivity to changes In _ and
NEM weight is shown in Table 2 for an aerobrake (at Mars capture and
Earth return) space vehicle configuration. An initial weight of
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1,434,215 POunds is required in low Earth orbit for the nominal case. If
the NN weight is increased by 15 percent, the initial weight in low Earth
orbit is increased by 6 percent over the nominal case. If the MEN weight
is increased by 15 percent, the initial weight is an increased by 4.2
percent over the nominal case.
The MEM initial weight sensitivity to variation in Nars lander
capsule and Mars ascent capsule weights is given in Figure 3. The
exchange factors are given in Table 3 for two different types of
propellants, N2 04/MNH and LOX/_gIH.
The initial mass required in low Earth orbit for each mission oppor-
tunity is given in Table 4 and Figures 4 and 5. The initial mass
required ranges from 1,280,001 to 3,575,321 lb for LOX/LH propellant.
These values do not include propellant boiloff in low Earth orbit during
orbital assembly time. The initial mass required in low Earth orbit for
the 1999 opposition outbound Venus swtngby using N 2 04 /NNH propellant is
8,869,090 lb. The initial mass in Earth orbit can be equated to cost and
used to determine the most favorable mission opportunities and the most
effective type of propellant for the propulsive stages.
PROPELLANT BOILOFF
The Mars mission is characterized by different mission environments
including LEO buildup, interplanetary transit, and Mars orbit. The
passive thermal protection on the cryogenic propellant tanks consists of 1
to 4 inches of MLI on the first stage and 4 inches MLI on the second and
third stage tanks. Vapor cooled shields are utilized on all tanks.
Table 5 relates cryogenic boiloff rate ranges for the different mission
environments using the all cryogenic vehicle configuration.
The boiloff rates for LEO were calculated with 1 and 4 inches of MLI
on the first stage. The boiloff rates in LEO are relatively high due to
large tank areas and albedo (reflected thermal energy from the Earth)
heating effects. The interplanetary transit mission phase Is character-
Ized by
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heating
minimize
Earth.
increased boiloff.
relatively low boiloff because of reduced vehicle tank area
two and three) and lower environmental heating. The lower
is contrlbutable to transit vehicle orientation during flight to
solar flux on tank wall areas and greater distance from the
Any deviation from the preferred orientation will result in
The Mars orbit stages experience medium boiloff rates
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through potential environmental heating due to vehicle orientations
driven by mission requirements. Preferred orientation in Mars orbit to
reduce environmental heating would lower the botloff rate.
1999 VENUS OUTBOUND SWINGBY MISSION ABORT
In the final selection of trajectories for the manned Mars stopover
missions, many factors other than vehicle weight must be considered.
Abort capability of the vehicle is one of these factors. It Is, there-
fore, necessary to plan and prepare for the possible irreparable failures
at some point during the mission.
Abort situations can be characterized as occuring In two different
phases of the mission which can be defined as (1) Earth departure phase
and (2) Heliocentric orbit phase. If abort maneuvers are executed within
30 minutes after trans-Venus injection, return to low Earth orbit can be
achieved within two days. The Interplanetary vehicle Is within Earth's
gravity sphere of activity up to 1 3/4 days after trans-Venus injection;
if abort maneuvers are undertaken within this time span, an elliptical
orbit return to low Earth orbit can be achieved within 18 days.
Heliocentric orbit phase is reached 1 3/4 days after trans-Venus
injection. The interplanetary vehicle (aero capture at Mars and aero
brake at Earth return) delta V capability after trans-Venus injection is
in excess of 9 km/sec for a small Earth return capsule; the Mars
excursion module has a 7.2 km/sec delta V capability and the second
stage main propulsion system has a 1.6 ka/sec delta V capability with
the total mission module weight of 113,633 lb. If mission abort is
executed sometlme less than 40 days into the mission, an Earth return
rendezvous trajectory can be achieved which returns back to low Earth
orbit within 80 to 250 days. After 180 days into the mission, the
Lnterplanetary vehicle Is committed to a Mars flyby which would return to
Earth in 560 days.
The above description of recovery from orbit conditions emphasized
mLnlmum delta V requirement for the return to Earth trajectory. Other
abort situations (i.e., abort after 40 days, braking into orbit at Mars,
Mars landing, Mars escape, etc.) need to be studied in more detail.
100
CONCLUSION
Comparative and sensitivity data have been developed for an opposi-
tion class Mars flyby and 60 day stopover missions to Mars. Also, data
were developed for conjunction class stopover missions. The 60 day
stopover mission utilized the.Venus swlngby mode in order to reduce the
propulsive energy required.
There Is a great variation in Initial mass required In low Earth
orbit for the all propulsive interplanetary space vehicles over a number
of misslon opportunltles. This varlatlon Is due to the eccentricity of
Mars orbit which has a perihelion distance of 1.38 A.U. and an apahellon
distance of 1.66 A.U. The wlde variation In initial mass may be reduced
by aerocapture at Mars and Earth return or by only returning to Earth
capture orbit wlth a small Earth return module and leaving the heavier
Mission Module in an Earth-Mars periodic orbit. The variation In initial
mass for the conjunction class mlsslon over a number of mlsslon oppor-
tunltles Is relatively small because there Is more freedom to optimize
the outbound transfer to Mars and thereturn transfer to Earth.
Mission abort capability, for the 1999 Venus outbound opportunity,
can extend out to 40 days after trans-Venus injection. In order to
minimize required weight in low Earth orbit, 4 inches of MLI on all
stages seems to be the most effective.
101
REFERENCES
i. Deerwester, Jerry M.: Inltlal Mass Savlngs Associated
wlth the Venus Swlngby Mode of Mars Round Trips. AIAA
2nd Aerospace Sciences Meeting, January 25, 1965.
2. Integrated Manned Interplanetary Spacecraft Concept
Deflnltlon r Vol. I, The Boeing Company, NASA Contract
NASI-6774, January 1968.
3. Soldner, John: Mars Mission Worklng Paper, Science
Application, Inc., April, 1985.
4. Young, Archle C., Mulqueen, John A., and Skinner, James
E.: Mars Exploration, NASA TM-86477, August, 1984.
5. Brown, Norman: Mars Mission Working Paper, Narshall
Space Flight Center, May, 1985.
102


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