The concept of the Mars airplane was developed as a potential vehicle for unmanned Mars exploration. It is suggested that its most appropriate use would be as an unmanned adjunct to a manned mission. Functions such as reconnaissance, exploration, remote delivery of instruments, etc., are possible. Several operational aspects of such a vehicle are unique compared to Earth operating aircraft

French, J. R.

English

NASA Technical Reports Server

N87-17753
THE I_RS AIRPLANE
J. R. French
Jet Propulsion Laboratory
Pasadena, CA
ABSTRACT
The concept of the Mars Airplane was developed as a potential
vehicle for unmanned Mars exploration. This paper suggests that its most
appropriate use would be as an unmanned adjunct to a manned mission.
Functions such as reconnaissance, exploration, remote delivery of Instru-
ments, etc. are possible. Several operational aspects of such a vehicle
are unique compared to Earth operating aircraft.
BACKGROUND
The Mars Airplane concept was developed by JPL and Dryden Flight
Research Center personnel in 1977/78 as a potential system for unmanned
exploration of Mars. The concept grew out of studies at Dryden of an
unmanned aircraft capable of operating for long periods at altitudes near
30.5 km (100,000 It). Such altitudes at low airspeeds dictated a non-
airbreathing engine. This together with the fact that atmospheric densi-
ty at 30.5 km on Earth is much like that near the Martian surface sug-
gested the possibility of a Mars airplane. Initial concepts by JPL and
Dryden looked promising and small contracts were let in 1977 and 1978 to
Developmental Sciences Inc. (DSI) to study Mars Airplane design (Ref. 1).
In addition, an ad-hoc science working group was convened to study what
the Mars Airplane might do (Ref.2).
This paper summarizes the results of those studies and adds some
additional thoughts of the author. Two versions of the airplane were
studied: 1) the cruiser, which could not land, but flew, taking data
until its fuel was exhausted and, 2) the lander which was capable of
repeated controlled landings and takeoffs. Only the latter version is
considered here since, for a manned mission, reusability of even robot
equipment is desirable.
ENGINEERING ASPECTS
Because of the limited data rate capability and desire to mlnlmtze
power requirements, the Mars Airplane was designed for a cruise speed of
90 m/sec (175 kt). The low atmospheric density (<1_ Earth) dictated an
airfoil with high cruise lift efficiency and a high aspect ratio wing.
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https://ntrs.nasa.gov/search.jsp?R=19870008320 2020-03-20T12:30:26+00:00Z
This combination together with the low gravity ylelds excellent range
capability as will be discussed later.
To minimize both mass and drag as well as for convenience in
stowage, an inverted Vee tall was used. The overall configuration
appears in Figure 1 (from Ref. 1). The resemblance to a modern high
performance sailplane is obvious. The dimensions are similar as well.
The propeller is quite large by Earth standards in order to perform
efficiently in the thin atmosphere. The tall is also large to allow for
large center-of-mass travel.
Martian conditions at these airspeeds result in a very low Reynolds
number for a vehicle of this size. Specifically, the Reynolds number for
the Mars Airplane at cruise is of the order of 4.5xi04 compared to 3x106
for a typical light aircraft. Most experience In thls range is with
free-fllght model aircraft. Thin high efficiency airfoils derived from
the work of Eppler at the University of Stuttgart were used in the DS!
design Ref.1. Since the time of that design (1978) there has been
increased interest in low Reynolds number airfoils and some of this new
work may be applicable.
The baseline powerplant for the Mars Airplane was a hydrazlne-fueled
reciprocating piston engine developed by Jim Akkerman of NASA Johnson
Space Center. This engine functions much like a reciprocating steam
engine. The hydrazine decomposes into a hot gas in a catalyst bed simi-
lar to that used in monopropellant rocket engines. This gas is then
valved into the cylinder of the reciprocating engine which vents it
overboard following the power stroke. This engine was flown successfully
on the Dryden "Mini-Sniffer" aircraft. The horsepower requirement for
the Mars Airplane as designed is 15HP. The complete engine when
developed for Mars should have a mass of about 13kg. The cruise specific
fuel consumption is expected to be 2.2 kg/HP-hr (4.85 lb/HP-hr).
An alternative powerplant was also investigated. This consisted of
a light-weight samarium-cobalt electric motor with a gearbox and a solid
state inverter. The unit was expected to weigh roughly the same as the
hydrazine engine, assuming very high performance of the Sa-Co motor. At
the time of the design, this was considered speculative by some and needs
to be reinvestigated. Electrical power was provided by primary
batteries, probably lithium thionyl chloride. The energy density usually
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quoted for these batteries is about 66 watt-hr/kg (300 watt-hr/Ib). For
this study a value of 649 watt/hr-kg (295 watt-hr/lb) was used. A much
higher energy density of 1199 watt-hr/kg (545 watt-hr/lb) was also
evaluated. This required heroic weight reduction measures in the battery
package and must be considered speculative.
Landing and takeoff were to be done vertically using variable thrust
monopropellant rocket engines derived from the Viking lander. The .fuel
for these engines is hydrazlne. In the case of the hydrazine engine
airplane, the rocket engines and the reciprocating engine would draw
propellant from a common supply, whereas the electric vehicle required a
hydrazlne supply just for the rockets. (Note that a disadvantage of the
electric airplane in regard to landing and takeoff is that it never
becomes lighter than it was InltlalIy, while the chemically powered ver-
sion loses mass as propellant is burned, thus making landing and takeoff
toward the end of a particular trip much less costly in fuel.) For a
manned mission, the airplane would presumably be reused in a manner
similar to remotely piloted vehicles (RPVs) on Earth. In this case, the
chemically fueled version may be more desirable, since rechargeable bat-
teries have much lower energy density than the primary batteries origin-
ally postulated. If the capabilities of the manned surface base included
In-Sltu Propellant Production (ISPP), the airplane could be supplied with
these propellants. Otherwise, residual propellant from the lander stare
should supply ample quantities. Engine types which might be considered
include reciprocating, gas turbine, and electric motors driven by fuel
ceils. (As an aside, if propellant manufacturing capability exists to
generate CO and 02 from the Martian atmosphere, a fuel ceil operating on
these materials would be most useful for airplanes, rovers and other
portable power needs rather than using the combination in a combustion
mode. This concept deserves further attention.)
PERFORMANCE
The hydrazine powered version of the Mars airplane was estimated to
have an operating ceiling of 15 km on a Mars standard day at minimum mass
(150 kg.) At maximum gross mass of 300 kg, the ceiling would be about 8
km, from which altitude the power-off glide range would be over 250 km.
Initial rate of climb at low altitudes is estimated to be 12.7 m/sec .P_
409
(2500 ft./min.), which is quite respectable and bodes well for terrain
avoidance capability and ability to cope with downdrafts.
Figure 2 (Ref. 1) presents range performance. The numbers are quite
respectable even if one ignores the rather debatable upper curve. These
numbers make no allowance for landing propellant, and would be typical of
a reconnaissance sortie with landing on a prepared surface at the base
rather than vertically. Maximum sortie radius would be half the range.
An early landing at high mass could reduce the range by 30-40_, while
retaining fuel for a final vertical landing at low mass would have a
lesser penalty.
OPERATIONS
The vehicle would be unstowed after landing and assembled by the
crew much in the manner of sailplanes on Earth. (This avoids one of the
serious drawbacks of the unmanned version; namely, how to achieve self-
deployment of a very complex folded structure while dangling from a
parachute.) Launch might be via catapult or the internal rockets.
Guidance and navigation would be preprogrammed, probably using lnerttal
systems with landmark identification for updates. Upon its return to the
vicinity of the base, control would be assumed by a crew member on the
ground who would control the landing, following normal Earth RPV
practlce.
Figure 3 shows a possible vertical takeoff profile suggested by DSI.
It is the author's opinion that propellant could be saved by starting the
propeller as soon as ground clearance is adequate and accelerating
directly into wing-supported fllght rather than following the lofted
trajectory shown. Landing would invoke a technique called "stable
stall".
Stable stall is a technique originally developed for the recovery of
free-flight model alrcraft. Briefly, it involves deflecting the horizon-
tal stabilizer to a very high angle, placing the aircraft in a deep
stall. The aircraft descends vertically in a flat attitude at a modest
and quite predictable rate. NASA studies show that the technique can be
satisfactorily applied to larger aircraft as well. In the case of the
Mars Airplane, the descent terminates in a rocket-braked landing. Figure
4 shows the profile of such a descent. Creation of a runway at the base
site would allow for conventional landing thus eliminating the need to
410
FI6, 2
CRUISE PERFORMANCE
HOURS OF
10000 "31"1 ENDURANCE
.--.....__ -9000 _,c "
_1_¢1_ (1799. tSj HOURSt
8000 __o, 1
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6000 _6.as
,,,5ooo__,
__4ooo
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¢: 20001 a
I000_
0
40 50 60 70 80 90 I00
P A Y L 0 A D (KG)
AUW 300 KG
WING SPAN 21 M
WING AREA 20 M z
CRUISE ALTITUDE I KM
LIFT/DRAG AT 300 KG "27.75
PROP EFFICIENCY .85
HYDRAZINE ENGINE
CRUISE SFC 4.85
ELEC. MOTOR EFFICIENCY .85
AUX. POWER CONSUMPTION .4 KW
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FIG. 3 BURNOUT
__ /jG30.5 m/sec
....
76.2m/see _ TRIM • X _. 1
(250ft/sec) _ _ -J_l_ _5 l
_.-% _ ' PULLOUT< L 5 G _, I
1(1000 ft ) . ..... # u_
VERTICAL TAKEOFF
FI6. q
1220 m
i (4000 ft) _-i
DRAGDECELLERATION
61 m/s
(200 ft/s)
61 m/s
(200 ft/s) DE-TRIM.
_OLLED
76 m/sec
(250 ft/sec)
A,I FALL LANDING 915m (31_ ft)305 000') m
m(1 _/(5 ft) 1.5
-t " ||
VERTICAL LANDING
412
carry propellant for a final vertical landing, on each sortie. To mini-
mize length of landing gear, all landings would be made with the propel-
ler stopped in the horizontal position. Mass could be saved by using
skids rather than wheels if only vertical or catapult takeoff is used.
FUNCTIONS OF A MARS AIRPLANE
An unmanned Mars Airplane could perform a variety of useful func-
tions in support of a manned Mars mission. Examples include:
I) Reconnaissance sorties to provide detailed route maps in
support of surface traverses by the crew in rover vehicles.
2) Scientific surveys of large regions or particular sites distant
from the base or otherwise difficult to reach.
3) Deployment of an array of remote observing stations
(penetrators, surface packages, or both) either by
air drop or by landing.
4) Delivery of high priority hardware to a crew far from base on a
surface sortie and/or return of priority samples, etc. from the
rover crew to the base.
Other functions will probably arise as the capability of the Mars
airplane becomes better understood and the mission definition improves.
Even the set of functions listed above could be of substantial benefit.
For example, detailed aerial maps will allow more rapid cross-country
traverses and warning of possible hazards. Optimum routes can be selec-
ted ahead of time. Large area aerial surveys supplementing work done
from orbit can be of great geologlcal significance. The abillty to
dellver science instrument packages to remote locations will be of sub-
stantial benefit, since it could not be done by surface rover but would
have to be done from space using individual entry packages. The ability
to deliver supplies of one sort or another to a rover crew could be vital
in case of hardware failure or a medlcal emergency.
MANNED MARS AIRPLANE
The question will inevitably arise as to feasibility of a manned
airplane for Mars. While (to the author's knowledge) this has not been
studied, there seem to be no technical reasons to prevent it. In fact,
the Mars Airplane described herein has adequate payload mass capacity to
carry any normal human. However, it could not carry a human in a space-
suit and full complement of llfe support equipment without exceeding the
413
design
payload.
larger than the vehicle described here, especially since
probably be desirable to carry a crew of two and a payload.
picture that develops is that of a vehicle of the general
appearance of the U-2, except being propeller driven.
The technical difficulties involved in creating such a vehlc]e do
not appear any more formidable than those involved in the small unmanned
craft. Some practical difficulties are stowage (especially for
atmospheric entry), assembly and handling on Mars, and propellant
consumption.
CONCLUSION
Based upon the work summarized in this paper, there appear to be no
serious technical difficulties involved in designing and operating a Mars
Airplane.
extending
safety.
REFERENCES
I •
2°
mass. Also, the compact fuselage volume lacks room for such a
A realistic manned Mars airplane would have to be considerably
it would
The mental
size and
It further appears that such a vehicle could be most useful in
the capability of the Mars surface crew and possibly enhancing
Development Sciences Inc., Final Technical Report on
JPL Contract No. 955012 "A Concept Study of a Remotely
Piloted Vehicle for Mars Exploration", August i, 1978.
JPL Publication 78-89 "Final Report of the Ad Hoc Mars
Airplane Science Working Group", November I, 1978.
414


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