In considering missions to Mars an important fundamental variable is the reduction in solar intensity at the li|1.6 million mile distance of Mars from the sun. ut this distance
Mars Mission
66
the solar intensity is reduced to a value of approximately 56 watts per square foot, which is U2% of the value found at the earth's orbital distance. Thus, if the Sunflower system were to be employed on a Mars mission, the solar collector size would have to be increased. Figure 16 has been prepared to determine Mars heat storage and percentage suntime operating requirements. This figure presents a plot of percentage sun-time, orbital period and absolute shadetime as a function of orbital altitude for circular Mars orbit in the Mars orbital plane. Reference to Figure 16 permits an evaluation of system requirements over any selected range of Mars orbital altitudes.
A system capable of operation at any Mars altitude from 1000 to 10,000 miles would weigh 962 lb and require a solar collec
tor diameter of feet.
Since it should be possible to establish both an earth departure and Mars arrival orbit sufficiently skewed to the ecliptic plane to avoid planetary shadow operation altogether, the weight of a system compatible with continuous sun opera
tion for a Mars mission is of interest. A Sunflower system can perform a Mars mission without planetary shadow require
ments by removing the lithium hydride heat storage capacity from the system and by sizing the collector area consistent with the reduced solar intensity at the Mars orbit. The Sun
flower solar collector area of 7Ul square feet may be reduced by its 62-1/2% overdesign factor which was required at the low altitude ecliptic plane circular earth orbit and subsequently increased by the factor 1/U2 to allow for the solar intensity at the Mars orbit. This gives a requirement for collector area of 1090 square feet and a collector diameter of 37·2 feet, to yield a total system weight of 681 l b . It is emphasized that this system weight synthesis is based on exactly the same performance assumptions which are being employed in the design of the current Sunflower system*
Venus Mission
Figure 17 presents the variation in percentage suntime, absolute shadetime and orbital period with orbital altitude for circular Venus orbits in the ecliptic plane.
In considering Venus missions it is observed that even for adversely shaded Venus orbits, the percentage suntime is never a problem since the solar collector must be sized for operating in the lower solar intensity environment of the earth's orbit. At its distance of 67·2 million miles from the sun, the solar intensity at Venus may be expected to reach 21$
watts per square foot, an increase of 91%. Therefore, the solar collector may be ignored in defining orbital limitations about Venus and it may be observed that the Sunflower system
S I X T H S Y M P O S I U M O N B A L L I S T I C M I S S I L E A N D A E R O S P A C E T E C H N O L O G Y
Fig. 16. Mars Orbital Characteristics.
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Fig. 17· Venus Orbital Characteristics.
S I X T H S Y M P O S I U M O N B A L L I S T I C M I S S I L E A N D A E R O S P A C E T E C H N O L O G Y
without modification will be capable of any orbit at altitudes below approximately 20.000 miles about Venus. Thus, the Sun
flower system need not be modified at all to accomplish the Venus mission and if earth departure orbits may be controlled
to exclude sunshade operation at departure, the resultant allowable reduction in collector size permits reduction of system weight to 630 lb while retaining 72-minute shade capa
bility at Venus. Such a 3-kw system would have a collector area of US6 square feet and a diameter of 25·9 feet.
Consistent with the assumption made in evaluating the Mars mission above, if sunshade operation can be avoided by sufficient departure from the ecliptic plane at arrival at Venus, the elimination of heat storage requirement allows re
duction in system weight to U50 lb.
Conclusions
Sunflower is applicable to a wide spectrum of missions both in terms of power level and system adaptation. The basic Sunflower concept is sufficiently flexible so that it may be packaged and utilized in several different configurations all of which use the components that are presently under develop
ment.
The wide applicability of Sunflower will cause it to emerge as one of a few standardized work-horse systems which will be used in sufficient numbers such that its inherent reli
ability goals are proved, Reference iw This benefit will be accompanied by a relatively low unit cost.
This paper has reviewed the factors which must be con
sidered when analyzing the use of Sunflower for a variety of mission categories. Even with the design flexibility inherent in the Sunflower system, it is apparent that the conventional approach to electrical power system integration must be modi
fied when considering Sunflower and other long duration, higher power systems. The effective utilization of these systems will require that they be given increased attention and that these operational requirements are fully considered early in the vehicle system design.
Finally, many applications are seen to be approaching within a time schedule such that Sunflower will be one of the few powerplants available in its performance class. For exam
ple, Dr. R. C. Seamans, Jr., associate administrator, NASA, testified recently that given sufficient budget, a man can be sent to the moon in six years, Reference 5 ·
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References
1. Daye, C.J., "A Family of Solar Space Power Systems Based on the Sunflower Concept,w 61-AV-31* presented at the Aviation Conference at Los Angeles. March 12-16, l?6l.
2. Rudy, J.A.> "Tne Sunflower Power Conversion System,11 13U9-6, presented at the ARS Space Power Systems Confer
ence, Santa Monica, California, September 27-30, I960·
3· Ross, i).P. and J.E. Taylor, ttTurboelectric Power Genera
tion for Space Vehicles,11 350A, 1961 National Aeronautic Meeting, Society of Automotive Engineers, ρ 2.
h* Cooley, W.C., " A Comparison of Nuclear and Solar Power Systems for Manned Space Stations,11 Proceedings of the Manned Space Stations Symposium, Los Angeles, California,
IAS-NASA-RAND, April 20-22, I960, ρ 212.
5# Aviation Week, April 21., 1961, ρ 31·