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R EARLY APPLICATION OF SOLAR-ELECTRIC n9 ^ FcF, w PROPULSION TO A 1-ASTRONOMICAL-UNIT 16^F^2 OUT-OF-THE-ECLIPTIC MISSION LLLOL
by William C. Strack and Frank J. Hrach Lewis Research Center Cleveland, Ohio
TECHNICAL PAPER proposed for presentation at Eighth Electric Propulsion Conference sponsored by W, the American Institute of Aeronautics and Astronautics Stanford, California, August 31-September 2, 1970 0
W
EARLY APPLICATION OF SOLAR-ELECTRIC PROPULSION TO A
1-ASTRONOMICAL-UNIT OUT-OF-THE-ECLIPTIC MISSION
by William C. Strack and Frank J. Hrach
Lewis Research Center Cleveland, Ohio
TECHNICAL PAPER proposed for presentation at Eighth Electric Propulsion Conference sponsored by the American Institute of Aeronautics and Astronautics Stanford, California, August 31-September 1, 1970
A NATIONAL AERONAUTICS AND SPACE ADMINISTRATION EARLY APPLICATION OF SOLAR-ELECTRIC PROPULSION TO A t-ASTRONOMICAL-UNIT OUT-OF-THE-ECLIPTIC MISSION
by William C. Strack and Frank J. Hra-h Lewis Research Center National Aeronautics and .race Administration Cleveland, Ohio
Abstract tational field of Jupiter to make plane changes. Large inclination angles to the ^^Jiptic are pos- Soiar-electric propulsion is evaluated for an sible if a close passage is made The advantage early application to an out-of-the ecliptic mis- of a Jupiter gravity turn, t:rwever, is tempered by sion. Relatively short flight times (100-475 days) increased mission time (it takes about 500 days are used to assess the performance of hardware gust to get to Jupiter) and th- restricted class of that could be built with present technology. The orbits that the spacecraft may attain after the electric prrpulsion system specific mass is as- Jupiter encounter. A particularly desirable mis- sumed to be 30 kilograms per kilowatt and current sion, for example, is to place a spacecraft in an thruster system efficiencies (e.g., 57 percent at inclined circular orbit at 1.0 AU, and this cannot 2600 seconds specific impulse) are employed. be reasonably done by means of a Jupiter swingby. Furthermore, the thrust program is simple - the This mission has several advantages that suit it thrust is constant and always directed normal to particularly well for early application: (1) data the instantaneous plane of the spacecraft orbit. are obtained at a constant 1 AU; thus, effects due The thrust is permitted to be turned off, however, to inclination are not obscured by effects due to and the typical trajectory is composed of several radius ; and can be compared more meaningfully with power-on and power-off constant-radius subares. existikq data, (2) Earth-to-vehicle communication Two currently available launch vehicles are as- distances are comparatively small, and (3) flu, ,o sumed: Atlas (SLV3C)/Centaur and Titan IIIC. the solar panels remains constant.
The results show that a negligible perfor- An alternative to using chemically-fueled mance loss is incurred by using the simple con- rocket propulsion to do this mission is to use low- stant-radius thrust control program compared with thrust electric propulsio ough analytical re- the more complicated (variable thrust direction sults have been generated t3 'S to suggest that an and solar power) variable-radius case. Also, a electric powered spacecraft can compare favorably fixed design spacecraft with 10 kilowatts of elec- with all-chemical propulsion systems - especially tric power and 2600 seconds specific impulse can at high inclination angles. The studies reported deliver nearly as much payload (never more than in Refs. 3 and 4 consider the case of constant 20 percent less) to a given heliographic inclina- power, variable radius trajectories that do not tion as an entire family of designs with optimum apply to solar-electric propulsion since solar values of power and specific impulse. This holds power varies with radius. Reference 5 gives re- true for both launch vehicles. sults for the case of thrust always directed normal to the instantaneous orbit plane in order to keep This fixed design electric spacecraft com- the spacecraft constrained to a constant radius of pares favorably with an uprated version (1040 kilo- 1 AU. This is a fairly simple thrust program to grams of propellant) of the Burner II chemical imr,lement and results in constant power output from stage. With the Atlas/Centaur, for example, the the solar panels. This avoids the problem of maximum '-.-liographic inclination attainable for matching a continuously varying power level to the 200 kilograms of net spacecraft mass is 25 degrees thruster system. However, the study was mainly for the uprated Burner II and 37 degrees for the limited to short time, single burn mission profiles fixed design electric spacecraft. With Titan IIIC using an Atlas/Centaur type launch vehicle. The these values are 27 and 41 degrees. In these present study generalizes this con.:ept to include s examples, the electric spacecraft requires a one- two and three burn mission profiles and also the year propulsion time, and about 470 days total to Titan IIIC launch vehicle. The objective is not reach maximum latitude compared to 91 days for the so much to present large amounts of parametric all-chemical systems. data, as in past studies, but to determine how well an early state-of-the-art solar-electric spacecraft Introduction would perform this mission compared to all-chemical systems, and to determine reasonable values of such The purpose of an out-of-the-ecliptic mission design variables as specific impulse, power loading, is to gather scientific data on interplanetary and propulsion duty cycle. fields and particles, ad to observe solar activity at high solar latitudes ?l) . All of such data ac- Analysis cumulated to date have been essentially within the ecliptic plane and primarily at 1 astronomical Trajectory Assumptions unit (AU). The past Mars and Venus probes have provided some limited data in the 0.7 to 1.5 AU Although most of the results given in previous range. Eventually this data base will be expanded studies are in terms of orbital inclination to the to include a wide range of distances from the sun ecliptic, it is considered preferable to present and inclination angles to the ecliptic plane. Near data in terms of inclination to the Sun's equator future plans, however, will be necessarily modest - since most of the interplanetary phenomena to bn especially in regard to the inclination angles due measured depend on heliographic coordinates. The to the very high energy expenditures normally re- ecliptic plane is inclined 7.2 0 to the Sun's equa- quired to make plane changes. One possible way to tor. Thus a launch to Earth escape velocity pro- avoid high energy expenditure is to use the gravi- duces an initial heliographic inclination i o of 1 TM X-52818
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1.26. Higher ent•rgy launches produce a hyperbolic inclination is not specified because, as will be excess velocity Vh that is most effective in seen later, optimal times exist within each trajec- changing inclination when applied at the nodes as tory class and these are determined by the optimi- sh .n, in sr.etch A. The electric propulsion system zation of the thruster shutdown and restart times. is assumed to be turned on soon after the high The electric power level is used in an iteration a thrust launch to at least escape energy; that is, loop to drive the final inclination to its _ desired in heliocentric space with velocity V. whose mag- value. The Lewis N-Body computer program l ' ) was nitude is identical to the Earth's velocity VEg used to calculate the trajectories and optimize in order to maintain a circular orbit at 1.0 Z. the free variables. From the sketch it is easy to show that the initial heliographic inclination is Chemical Systems Assumptions
Vh The assumed launch vehicle performance is 1 = ;.2 0 + 2 sin -1 (1) shown in Fig. 1 for the Atlas/Centaur and the 2 V° Titan IIIC. The launch mess against burn-out velocity V at 185 kilometers altitude comes from The Earth's gravitational effect is ignored Ref. 8. The hyperbolic excess velocity Vh is when the spacecraft is in heliocentric space. The determined by the booster burnout velocity V thrust is directed normal to the instantaneous and circular orbit ve'.ccity Vc: orbit plane so that the size and shape of the orbit are not changed as the inclination is increased. Vh=Vb -2V2_(4) This could be done without thruster gimbaling by orienting the thrust vector with the spacecraft Equations (1) and (4) are combined with the curves attitude control system. Regardless of how the iii Fig. 1 to yield the relationship between launch thrust vector is controlled, at least some attitude mass and initial inclination. control is required to keep the solar panels facing the Sun. The time rate of change in inclination is Also shown in Fig. 1 is the performance data for these two bgpsters with an uprated Burner II i = .^ cos u (2) stage added on l J . The 1.040 kilogram propellant d 0 loading version of$finer II assumed here is not where a is the thrust acceleration and u is the currently available Nevertheless, it is used argument of latl.ude (sketch b). The rate de- in the comparison with the solar-electric system creases as the spacecraft moves away from the node in order to compare both system types at approxi- and is zero at the first antinode (u = 1/2) -- a mately the same technology level.. position of maximum distance from the Sun's equa- torial plane. Beyond the first antinode the thrust Solar-Electric Spacecraft Assumptions direction must be reversed in order to continue increasing orbital inclination. And again when One electric propulsion system specific mass the spacecraft reaches the second antinode a is assumed to be 30 kilograms per kilowatt. (u = 3n/2), the rate of change of inclination This number is generally considered to represent vanishes and the thrust direction should be rever- hypothetical solar-cell powt,igd is cecraft at the sed. IIence, even for this simplified thrusting present level of technology The propul- method, continuous control Df the spacecraft atti- sion system mass' includes both the power and tude is required and occasional complete reversals thrust subsystems defined by the suggested of thrust direction are necessary. The thrusting nomenclature in Ref. 12. The power subsystem in- program is illustrated in si;etch b for a three-burn cludes primary power, thermal control, cabling, class trajectory, although one- and two-burn tra- support structure, etc. and the thruster subsystem jectories are also included In the study. In the includes thrusters, power conditioning control, majority of cases considered the thrusters are cabling, support structure, etc. The tankage mass turned off near the antinodes because the ineffec- mt is assumed to be 10 percent of the propellant tiveness of thrusting there results in a payload mass T°p and includes tank structure, plumbing, penalty, as will be shown later. residuals, reserves, etc. With these assumptions the net spacecraft mass mn is Thrusting also changes the line of nodes (sketch b) while the orbit inclination is increas- ing. mn - mo - mps - mp - mt (5)
The rate of change in the longitude c,f ascend- = m° - aPe - mp - 0.1 mp (6) ing node is The net spacecraft mass includes more than a sin u (3) just the scientific payload. It also includes T.-Tn—i equipment for guidance, attitude control, thermal The change in n between the first and third control, support structures, communications, data nodes is generally 10 to 35 degrees forward. handling and computation. In these expressions Equations (2) and (3) as well as the other equa- m° is the initial spacecraft mass and P e is the tions of motion are derived in Ref. 6. required electrical power. P e is calculated from the useful kinetic power in the jet exhaust Pj Unless otherwise specified, the net spacecraft and thrust subsystem efficiency Tits' mass is maximized for a given final heliographic inclination by optimizing: the launch energy 1 V2 2 o ), the electric thruster specific (equivalent tn i P = P.2^j=mgols impulse, and all the thruster shutdown and restart e times. The total time required to achieve a given i ts its 2^ts
2 I®R t Y
whe-e V is the exhaust velocity, I s 1s the (1) those trajectories that have a single thrust spectfic^impulse, and g is the gravitational arc, (2) those that have one coast arc between two constant ( g - 9.80665 m9s 2 ). The thrust subsystem thrust arcs, and (3) those that have two coast arcs efficienc,/y o ^ ts is the product of the thruster and three thrust arcs. Additional flight time efficiency and theower conditioning effi(,iency would of course result in additional trajectory (azsumed to be 0.84: classes involving even more coast and thrust arcs. For conciseness the classes considered herein are 0.88 Eo simply referred to as single-burn, two-burn, and (e) three-burn trajectories. The circled points in T1 ts = ^10 Fig. 2 show the performance increase due to re- 2 1 laxing the no'coast constraint at three specific I, flight times. There is no benefit with the single- burn class since it is identical with the all pro- This equation is based on an idealized thr}u^ t^r pulsion class. The two-burn benefit is a 5 percent (aseuming constant ionization power losses 13 ) increase in net spacecraft mass and the three-burn and has been found to correlate experimental re- benefit is 6 percent. The power levels (not shown) sults of real thrusters reasonatly well. E is are essentially unchanged. Thus the performance the asymptotic vali!.e of thruster efficiency oat advantage of coasting trajectories is not great. infinite Is and Io is the specific impulse for However, the thrust direction must be changed by a thruster efficiency of 1/2 E F,q. (8) fits 180 degrees at the antinodes regardless of whether the curve in Ref. 14 labeled "Kture 2-3 kVt' by coast arcs are permitted or not, and it might be letting Eo = 0.85 and 1 = 1465 seconds. In necessary (although unlikely) to shut the thrusters view of current data for 30 centimeter diameter down during thir reorientation maneuver to avoid insulated grid thrusters, :Ref. 14 suggests regard- disturbance torques. Also, scientific data gather- ing this curve as 1968 "pr esent" curve at the 2 to ing and communication is most desirable at such 3 kilowatt level. Reference 3 predicts the mid- times and the extra power made available by thrus- 1970's technology level with an efficiency curve ter shutdown might be used for these other purposes that is represented by Eq. (8) by setting E o = For all of these reasons, the remaining data will 0.957 and Io = 1630 seconds. To be conservative be shown only for the optimum flight times with i most of the calculations em ploy the 1968 "present coast arcs permitted. data but for the sake of comparison some projected mid-1970's data is also used. Performance of Electric and Chemically Powered Spacecraft Results and Discussion The potential performance of a solar-electric Flight Time and Trajectory Classes system is compared to the all-chemical systems in Fig. 3. Part a is for the Atlas/Centaur while Typical results of net spacecra_t mass as a part b is for the Titan IIIC launch vehicle. Net function of flight time are presen , red in Fig. 2. spacecraft mass is plotted as a function of final The particular system illustrated ::s a Titan IIIC- heliographic inclination for the launch vehicle launched solar-electric spacecraft that attains a by itself (dotted curve), the launch vehicle with heliographic inclination of 30 degrees. The solid the uprated version of the Burner II stage added curve represents the restricted case of all pro- (solid curve), and the launch vehicle with fully pulsion -- no coast subares are permitted. Gen- optimized solar-electric spacecraft (the space- erally, this curve rises rather steeply and shows craft iesign change;, along each curve) added the marked payload improvement possible with in- (dasi l ed curves). There are three curves for the creased flight time. The flight times in Ref. 5 electric spacecraft -- one for 1-burn trajectories, are primarily in the 80 to 100 day range which, 'n one for 2-burn trajectories, and one for 3-burn this case, are not attractive -- yielding only 3b trajectories. The time required to reach an tLnti- kilograms of net spacecraft mass. But increasing node following the final thruster shutdown is also the flight time to 275 days raises the net mass to noted for each curve. This time is treated as the 44C kilograms while 440 day trips provide a further mission time (i.e., rather than the time to attain increase to nearly 700 kilograms. a specified inclination) since the scientific data to be collected is of most interest at the anti- nodes. Actually, the mission time for the solar- Note that three distinct local maxima exist electric case varies slightly with inclination, that are some six months apart. These occur be- but since the variation is only several days an cause the spacecraft is constrained to a continuous average time is quoted. thrust program that is relatively ineffective every six months when the craft is near an antinode (see The values of net spacecraft mass of main 4 Eq. (2)). For example, if the powered flight time interest lie approximately between 200 and 400 is 275 days thrusting terminates 22 days before the kilograms. These estimates come from related second antinode is reached; but for 300 day flights mission spacecraft such as the 400 kilogram the thrusters operate 7 days after the second anti- Mariner 7 and the proposed 210 kilogramsp cecraft node. Hence the extra 25 days of the 300 day mis- for project HELI05. The HELI06 mission( 15^ is a sion are spent wastefully by thrusting in the prox- 0.3 AU solar probe with some 50 kilograms of imity of an antinode. The net result of this in- scientific experiments aboard. Both launch vehi- efficiency is an 8 percent drop in net spacecraft cles are limited to heliographic inclinations of mass compared to the 275 day flight. 19 degrees if no upper stage propulsion is used. Adding an uprated Burner II stage would raise If the no-coast constraint is removed, the this limit to about 25 7 ,grees for the Atlas existence of these three local maxima result in Centaur or 27 degrees for the Titan IIIC for 200 the definition of three distinct trajectory classes: kilograms of net spacecraft mass. If hypothetical and fully optimized solar-electric systems are valuer. The sensitivity data are given with the substituted for the Burner II, the performance underlying idea of fixing the spacecraft design. is improved only if 2- or 3-burn trajectories are used. The 2 -burn trajectories would allow 32 Electric Power Requirements degrees for 200 kilograms net spacecraft mass us- ing the Atlas/Centaur or 36 degrees using the In Fig. 5 the optimum electric power level Titan IIIC. The 3-burn trajectories would permit is plotted for both launch vehicles as a function 39 degrees using the Atlas /Centaur or 43 degrees of final heliographic inclination. The Inclina- with the Titan IIIC (not shown). Roughly speaking tion values that correspond to 200 and 400 kilo- then, an uprated Burner II looks attractive in the grams of net spacecraft mass (from Fig. 3) are 20-25 degree range while solar-electric spacecraft noted on each curve. These net mass values bracket look attractive in the 25-43 deg:ree range. the range of primary interest and together with 2- and 3-burn trajectories lead to optimum power It must be emphasized that the solar-electric levels that are surprisingly constant. For exam- data in Fig. 3 re presents a whole family of space- ple, the best power using the Atlas/Centaur varies craft, optimized with regard to specific impulse, only between 10 and 11 kilowatts for 200 to 400 installed electric power, launch velocity, and kilograms o' net spacecraft mass. This occurs coast arc timing. The data, therefore, do not between 26 and 32 degrees for 2-burn trajectories reflect the performance of a single spacecraft and between 31 and 39 degrees for 3-burn trajec- design. Such a single design would have fixed tories (from Fig. 3). The optimum power using values of specific impulse and electric power, the Titan IIIC varies between 16 and 20 kilowatts although launch velocity and coast arc timing for the same spacecraft size range. Thus, the could still be optimized. The actual performance optimum electric power for the Titan IIIC launched of such a single design is presented later in this spacecraft is 1.7 times the optimum power for the report and affords a fairer comparison of electric Atlas/Centaur launched spacecraft -- roughly the and chemical propulsion. same ratio as the launch vehicle capabilities near escape speed. The improved performance of the 2- and 3-burn solar-electric propulsion systems comes at the These power levels are optimum in regard to price of increased mission time. Actually this payload capability only. Since solar-electric penalty is not overbearing since for this mission spacecraft are relatively expensi-e (e.g.,, Hi.con the spacecraft can gather important data all along solar cell arrays cost about $300 per watt its transfer trajectory. This is illustrated in the complete system is likely to be more cost Fig. 4 where the distance from the Sun's equa- effective at reduced power levels if the associated torial plane is plotted against time for a 200 payload penalty is not too large. The lower part kilogram spacecraft launched by Atlas/Centaur. of Fig. 6 shows two typical tradeoff curves of The all-chemical system rises continuously to net spacecraft mass against installed electric reach its maximum distance of 0.42 AU in 91 days. power. It is immediately apparent that the elec- The distance from the Sun's equatorial plane would tric spacecraft performance is attractive over e. then continue in a sine-wave pattern, reaching rather broad range of power level. Consider first 0.42 AU below the equatorial plane six months the Titan IIIC launched spacecraft mission to 40 later and then returning to the maximum latitude degrees heliographic inclination with 3-burn class point above the plane six months after that. trajectories. The optimum power level of 19 kilo- watts yields nearly 300 kilograms. At 15 kilo- The electric spacecraft generates a pattern watts the net spacecraft mass drops 4 percent to similar to a sine-wave but with increasing ampli- 285 kilograms, and at 10 kilowatts it drops tude during the propulsive pEriods. It reaches 19 percent to 240 kilograms. This figure also 0.31 AU at the first antinode in 91 days, 0.46 AU shows that the optimum specific impulse decreases below the equatorial plane six months later, and from 3400 to 2600 seconds when the power is re- 0.63 AU above the plane six months after that. duced from 19 to 10 kilowatts. And the optimum It would reach the all-chemical limit of 0.42 AU initial spacecraft mass decreases 33 percent - in 261 days. The total electric propulsion time from 1660 to 1120 kilograms. The specific impulse is 365 days (8800 hours) which is also a typical variation is not particularly important, but the thruster lifetime estimate for near-term mission 45 percent reduction in electric power and the applications. It is also important to note that 33 percent reduction in initial spacecraft mass the 25 degree heliographic inclination achieved might very well be worth the 19 percent net space- by the all-chemical spacecraft would be attained craft mass penalty. by the electric spacecraft if its propulsion system functioned for only 167 days (4500 hours) Similar tradeoff results are obtained for of operation. Thus, since t the time of this Atlas/Centaur launched spacecraft to 30 degrees writing the SERT II mission has already demon- using 2-burn trajectories. However, in view of strated flight-rated thruster subsystem lifetimes the fact that 10 kilowatts is about optimum for of four months, one can be reasonably certain that this launch vehicle, as well as being a good com- even a near term electrically propelled spacecraft promise choice for the Titan IIIC, it might be wise would succeed in reaching at least the all-chemi- to consider a standard 10 kilowatt design that cal propulsion inclination limit, if not consider- would nicely match both boosters. Furthermore, ably more. this idea is reinforced by the previous result that the range of optimum power levels is quite All of the data shown in Figs. 3 and 4 repre- broad. This implies that the 10 kilowatt power sent systems optimized to deliver maximum payload. level is a reasonably good choice over the entire The corresponding values of the electric power spectrum of interesting missions (defined earlier level and thruster specific impulse are presented as those missions that can be accomplished with next along with the effect of using nonoptimum 200 to 400 kg of net spacecraft mass). More evi-
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deuce for this conclusion will be given after a step toward the evolution of a multimission solar- good compromise value of the specific Impulse is electric spacecraft. Some a-Alysis work has al- also obtained. ready been done in this area. In Ref. 18, for example, a 6.5 kilowatt, 3500 second specific im- Specific :mpulse Requirements pulse design is analyzed In depth for four differ- ent missions. From Ref. 19 a 10 kilowatt, 3250 The optimum specific Impulse values are given second specific impulse design appears to be at- In Fig. 7. The results are very nffarly independent tractive for four missions. It is clear that if a of the launch vehicle and are therefore plotted as number of different missions are considered the a ;Ingle set of curves. The average value in the power and specific impulse values found to be 30 to 40 degree range is 2900 seconds for 2-burn attractive here for the extraeeliptic mission would trajectories and 3650 seconds for 3 -burn tra„ec- probably be replaced with a new set that would be tories. As in the case of electric power, however, an appropriate compromise for many dissimilar mis- net spacecraft mass is not particularly sensitive sions. Further work in this area is recommended to specific impulse az is shown in Fig. 6_ For the and should include a wider range of missions and same two example missions discussed previously, launch vehicles. sp e cific impulse values br!tween 2500 and 4700 sec- onds result in no more than a 15 percent penalty State-of-the-Art Effects In net spacecraft mass. The figure also shows that significantly less electr'^ power is required if Electric propulsion assumptions. - All of the the specific impulse is lowered from its optimum preceding results are for state-of-the-art inputs value. In order to pick a good compromise value of assumed to be current. The complete propulsion specific impulse for a fixed spacecraft design, system specific mass is assumed to be 30 kilograms note first that if the 10 kW constraint is imposed per kilowatt, the tankage fraction is 10 percent, the performance of the Titan IIIC case is affected the thruster efficiency curve reflects current much more than that of the Atlas/Centaur. There- designs, and a simple, nonoptimum thrust control fore, using the optimum specific impulse for the is employed. The effect of altering thes, parti- 10 kW Titan IIIC case would result in a good over- cular state-of-the-art assumptions is shown in all compromise provided that the performance of the Fig. 11 for a 3-burn Atlas/Centaur launched space- Atlas/Centaur is not seriously affected. From craft. The solid curve repeats earlier data and Fig. 6, this value of specific impulse is 2600 reflects the current state-of-the-art assuunptions seconds. Figure 8 shows that for this specific lust specified. All other curves are to be com- impulse the optimum power level for the Atlas/Cen- pared to it. The dashed curves f'or estimated taur case is A.7 kW -- which is close enough to mid-1970's technology thrusters (3) -- about 6 the 10 kW constraint value to suggest that 2600 percent more gfficient -- and 3 percent tankage. seconds is indeed a good compromise value for all The performance gain is generally rather modest -- cases of major interest. allowing an additional 2 degrees of inclination, for example, at the net spacecraft size of 200 A Fixed Spacecraft Design kilograms. The two dotted curves show the effect of assuming the propulsion system specific mass to The results of the previous two sections be 25 and 35 kilograms per kilowatt instead of suggest that the use of 10 kilowatts of electric 30 kilograms per kilowatt. Again the performance power and 2600 seconds specific impulse for a fixed change is not large -- between one and two degrees design spacecraft might result in a good overall of inclination. The predicted mid-1970's techno- power and payload tradeoff over the whole mission logy gain would offset a specific mass increase of spectrum of interest. That this is indeed true is around 5 kilograms per kilowatt. On the other hand, shown in Fig. 9 where both the family of oy)timum gains from both improved thrusters and decreased designs and the single fixed design are compared. specific mass are additive. The fixed design performance penalty ranges from negligible to a maximum of 20 percent at 45 degrees Optimal thrust control. - Several sample cases using the Titan II1C. of optimal thrust control were generated with the computer code described in Ref. 20. In this case The performance of this single electric space- the spacecraft can no longer be constrained to a craft design is compared to all-chemical systems radius of 1.0 AU because the thrust is not required in Fig. 10. This figure is the same as Fig. 3 ex- to be normal to the orbit plane. The payload gains cept that it concerns one spacecraft design instead over the fixed thrust program are so negligible of a family of optimum designs. Comparing; these (e.g., less than 1 percent) that a comparison curve two figures shows that imposing the single design could not be drawn on Fig. 11 distinct from the constraint on the electric system does not mater- reference curve. This is a particularly important ially affect the comparison with all-chemical sys- result for early application missions since the tems. The single design electric propulsion system possible penalties for strictly optimal thrust pro- can deliver far more net spacecraft mass than the gramming (increased subsystem weight, cost, and uprated Burner II and also extends the maximum in- less reliability) are avoided. For solar-electric clination limit (for 200 kg of net mass) from 25 to propulsion systems, the optimal trajectories are 37 degrees using Atlas/Centaur and 27 to 41 degrees almost identical with +.he constrained trajectories using Titan IIIC. the spacecraft stays at essentially 1.0 AU rather than drifting outward to Maro orbit as in the case Whether or not one would actually use a single of the constant power trajectories reported in fixed design electric spacecraft for various mis- Refs. 3 and 4. This difference arises because as sions remains an open question. What is illustra- the distance from the Sun increases the solar panel ted here is that if such a spacecraft did exist it output decreases in approximately an inverse square would be quite versatile indeed. Extending this relationship. concept to include completely different missions (e.g., close solar probes) is the next logical
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Launch vehicle performance. - The Atlas disadvantages of higher space vehicle cost and (SLV3C entaur and Titan IIIC boosters are assumed longer flight time is not dealt with here. To in thi:: otudy because they already exist and will avoid costly development of systems that are best prub;sb.ly be available in the mid-1970's. At the suited to isolated cases only, considerations such time of this writing, the only booster larger than as these should, in fact, be viewed from an overall these (excluding Zaturn V) that seems certain to space program standpoint rather than for a single exist by 1975 is the unbuilt Titan IIID/Centaur. mission. It is clear, nonetheless, that a 1 AU It is the planned launch vehicle for the 1.,75 out-of-the-ecliptic mission is particularly well- Viking misstep to Mars and has considerably higher suited to solar-electric propulsion and the re- performance 1 than either the Atlas/Centaur or latively simple normal-to-the-orbit thrust control Titan IIIC. The performance growth potential using requirement enhances its prospects for early appli- the Titan IIID/Centaur for the out-of-the-eclipti cation. Simplified navigation and trajectory re- ..fission is illustrated in the table below. The quirements are other desirable features of this olar-electric data is for the 10 kilowatt, 2600- mission. Considering all of these factors together econd specific impulse fixed design spacecraft. leads to the suggestion of doing this mission with a first generation solar-electric spacecraft in System Heliographic Inclination order to flight-test new hardware as well as col- for 200 kg of Net lect scientific information. Spacecraft Mass, degrees References Atlas/Centaur 19 Atlas/Centaur/ 25 1. Biermanr,, L., "Some Asp•cts of the Physics of uprated BII Interplanetary Space belated to Out-of-the- Atlas/Centaur/ 37 Ecliptic Studies," Advances in Space Science solar-electric" and Technology, Vol. 7, F. I. Ordway, III, ed., Academic, New York, 1965, pp. 437-447. Titan IIIC 19 Titan IIIC 27 2. Minovitch, M. A., "Utilizing Large Planetary uprated BII Perturbations for the Design of Deep-Space, Titan IIIC 41 Solar-Probe, and Out-of-Ecliptic Trajec- solar-electric* tories," TR 32-84P, NASA CR-69222, Dec. 15, 1`165, Jet Propulsion Lab., California Inst. Titan IIID/Centaur 29 Tech., Pasadena, Calif. Titar IIID/centaur/ 34 up'. -ated BII Mascy, A. C., "Extraecliotic 1.0 a.u. Constant Tit•.n IIID/Centaur/ 51 Power Electric Propulsion Mission," Journal r
Concluding Remarks 8. Anon., "Launch Vehicle Estimating Factors," NASA Office of Space Sciences Applications, What is shown in this study is that current Washington, D. C., Jan. 1970. electric propulsion technology could produce a spacecraft that Ields important performance ad- 9. Anon., "Mission Planner's Guide to the Burner vantages compared to all-chemical systems. Hcw II," D2-82601-5, Apr. 1968, Boeing Co., one weighs the advantages of higher performance Seattle, Wash. and smaller launch vehicle requirements with the 10. Ratcheson, W. I., I'Fabrication FeFsibility Study of a 20 Watt per Pound Solar Cell Array," D2-23942-5, NASA CR-70582, Nov. 19, Using 3-burn class low-thrust trajectories 1965, Boeing Co., Seattle, Wash. (average propulsion time is 360 days, average time to attain final inclination is 442 days, and average mission time is 467 days.)
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11. Stager, D. N. and Anderson P. N., "An 850 16. Kerslake, W. R., Byers, D. C., and Staggs, Pound, 20 kW Solar Arrays Paper 65-471, J. F., "SERT II: Mission and Experiments," July 1965, AIAA, New York, N. Y. Journal of Spacecraft and Rockets, "al. 7, No. 1, Jan. 197C, pp. 4-6. 12. Anon., "Eaectrlc Propulsion Mission Analysis: Terminology and Nomenclature," SP-210, 1969, 17. Boretz, J. E., 'Large Space Station Power NASA, Washington, D. C. Systems," Paper 68-1034, Oct. 1968, AIAA, New York, N. Y. 13. Kaufman, H. R., "The Electron-Bombardment Ion Rocket," Advanced Propulsion Concepts, Vol. 1, 18. Anon., "Study of a Solar Electric Multi-Mission Gordon and Breach, New York, 1363, pp. 5-18. Spacecraft, , Rep. 09451-6001-RO-02, Jan. 15, 1970, TRW, Inc., Redondo Beach, Calif. (Work 14. Richley, E. A. and Kerslake, W. R., "Bombard- under contract JPL-952394.) ment Thruster investigations at the Lewis Re3earch Center," Paper 68-542, June 1968, 19. Nagorski, R. P., "A Solar Electric Spacecraft AIAA, New York, N. Y. Tr-jectory and Performance Analysis for Area Target Missions," Paper 70-212, ,Tan. 15. Ruppe, H. 0., "Solar R-search in the Space 1970, AIAA, New York, N. Y. Age," Paper presented at the United Nations Conference on Peaceful Uses of Outer Space, 20. Hahn, D. W., Joh.-ison, F. T., and Itzen, B. F., Vienna, Austria, Aug. 1`.768 (rev. June 1969). 11 Chebychev TrLjec„ory Optimization Program (CHEBYTOP)," D2-1213CB-1, NASA CR-73351), July 1969, Boeing Co., Seattle, Wash.
7 4000
M O r- 2000 W 1000 rn Y 800 E 600 V TITAN IIICIBII (1040) 400
U 200 TITAN IIIC
ATLAS (SLV3C)1 100 CENTAUR 80 ATLAS (SLV3C)/CENTAUR113II (10401 60 5011 - I I I l 12 13 14 15 16 17 RURNOiJT VELOCITY AT 185 km ALTITUDE, V b, kmis ll I I I I i 0 2 4 6 8 10 12 HYBERBOLIC EXCESS VELOCITY AT 185 km ALTITUDE, V h, kmis Figure 1. - Launch vehicle performance.
'41, y
6
Vo Vh ECLIPTIC PLANE
U. ^u VL NODE o ^^' 2 —SUN'S EQUATORIAL w PLA NE • (a)
FIRST ANTINODE-,
SECOND THIRD ANTiNODE—_,^^ i NODE SUN'S EQUATORIAL PLANE--\
RADIUS
^– SUNS `-THRUST OFF FIRST NODE io \THRUST i THIRD NODE ON LSECOND ANTINODE NOTE: ARROWS DENOTE THRUST DIRECTION (NORMAL TO ORBIT PLANE)
(b) 6
POWER-COAST-POWER COAST-POWER 13 BURNSI,- CY 700 tr) 0 fl- LO w 600
Y POWEP E 500 POWER ,„ V)
L.L.~ 400 Q 0. U w U Q 0. N 300 z
200
POWER 100 ONLY I1 BURNI A-.,• 0 100 200 300 400 500 TIME TO ATTAIN 30 DEGREES HELIOGRAPHIC INCLINATION, DAYS Figure 2. - Effect of flight time on performance of solar- electric spacecraft using Titan III C launch vehicle. Also shows effect of all-propulsion constraint. System variables optimized. BOOSTER ONLY BOOSTER PLUS UPRATED BURNER II 1200 ` -- BOOSTER PLUS SOLAR-ELECTRIC
M 800 r-O LO ^- 3-BURN 1 BURN - w \ ^^ ,^^ ^2-BURN MISSION ` o, TIME, DAYS \ 467 91 11U^^ 1 p —L_^ 288 E IAI ATLASICENTAUR. V) 2000
Q ^ CK 1600 C-) Q ^ a \ \ WzF- 1 200 `\ \\ ,^ 3-BURN 800 — 2-BURN
400 MISSION \^` ^` `\ 467 TIME, DAYS 1-BURN \ 91 91 p 110 288
10 15 20 25 30 35 40 HELIOGRAPHIC INCLINATION, DEG
IBI TITAN IIIC. Figure 3. - Potential performance of solar-electric space- craft compared to all-chemical stages. Out-of-the- ecliptic mission. Burner II propellant loading, 1040 kg; electric system total specific mass, 30 kglkW. All free variables optimized.
6 0 1
— t — HELIOGRAPHIC 8 COAST FLIGHT POWERED FLIGHT INCLINATION Wz EQUALS 39° Ii i 0 .6 ATLASICENTAURI i I, UPRATED BURNER II ° r HELIOGRAPHIC INCLINATION —^ Cr ATLAS/CENTAUR/ .4 +^ /* EQUALS 250^ SOLAR-ELECTRIC Mo v,
LO Z % I 1 W N I .2 LL. / W U ^ Z ^ Q N I ° 0 100 200 300 400 500 TIME, DAYS Figure 4. -Time history of distance from Sun's equatorial plane. Net spacecraft mass, 200 kg. All free variables optimized.
O 400 kg OF NET SPACECRAFT AMASS 25 p 200 kg OF NET SPACECRAFT MASS
TITAN IIIC V Y --- — — ATLAS/CENTAUR NUMBER 20 CL OF BURNS
W 3 0 15 a _U \2 WU' 10 J W
^^ 1 \2 5 a 0
01 11 1 1 1 1 20 25 30 35 40 45 HELIOGRAPHIC INCLINATION, DEG Figure 5. - The optimum electric power requirements for out-of -the -ecliptic missions using electric propulsion. All free variables optimized. I
2500 QU Y °- o TITAN IIIC TO 400 `^ E IN 3 BURNS aJ ^; = N 1500 V) ►z ^ 0 L ATLASICENTAUR TO 30° IN 2 BURNS w ^Q 500 o
4000 U ATLAYCENTAUR TO 30° ^i N U IN 2 BURNS w in
N 3000 TITAN III C TO 400 ^ J IN 3 BURNS
a 0 2000
320 x m . y Y EC
v 280 TITAN IIIC TO ATLAS/CENTAUR 40" IN 3 BURNS l"Z TO 30° Le IN 2 BURNS WU 240 Q NCL 11 Z ^L / I / I I I I 0 5 10 15 20 25 ELECTRIC POWER, P e, kW Figure 6. - The effect of using nonoptimum electric power level for out-of-the ecliptic missions. 0
O r- N 5000 _CA W W N J ZD Cd 'l000 C
U
U w a V) 3000
CL O 2000 20 25 30 35 40 HELIGUG IR AI HiC, INCLINATION, DEG Figure 7. - The optimum specific impulse requirements for out-of-the-ecliptic missions using electric propul- sion. , Atlas/Ce,itaur and Titan IIIC launch vehicles. All free variables optimized.
r, 6
25
Y a 20 TITAN III C TO M 400 IN 3 BURNS O w O W d 15
wU w 10 ATLAS-CENTAUR TO 30° IN 2 BURNS a O 51
o, 320 Y
L" E TITAN III C TO cn 280 40° IN 3 BURNS
tz a W 240 U ATLAS/CENTAUR TO 300 IN 2 BURNS Ln w Z 200 2000 3000 4000 5000 6000 SPECIFIC IMPULSE, I s, s Figure 8. - The effect of using nonoptimum specific impulse for out-of-thP-ecliptic mis- sions.
0
OPTIMUM DESIGN 1000 — — — — FIXED DESIGN 0 f– TITAN IIIC IN 2 EURNS w Y 800 .,
c ^ E ^r-TITAN IIIC IN 3 BURNS
V1 600
\ \\
ATLAS/CENTAUR^`` ^^ \^ \ \ ` ^^ a IN 3 BURNS
Uj —71 200
ATLAS/CENTAUR IN 2 BURNS-A, 0 1 1 1 _ 1 25 30 35 40 45 HELIOGRAPHIC INCLINATION, DIG Figure 9. - Comparison of fixed spacecraft design with family of optimum designs for out-of -the -ecliptic missions. Fixed design power level, 10 kW; specific impulse, 2600 seconds.
•
— — — — BOOSTER ONLY BOOSTER PLUS UPRATED BURNER II 1200 -- BOOSTER PLUS SOLAR-EI.ECTRIC
M O 800 ^\ r- i r 3-BURN e W 2-BURN 400 MISSION \` 1`^^^9 467 Cn Y TIME, DAYS 91 1 ^ 288 C 0 E (A) ATLAS/CENTAUR. V ^n 2000
1600 \\ \ wU U Q LL 1200 V)
z `^ r 3-BURN 2-BURN
400 MISSION \\ TIME, DAYS X 91 ~^~^ 467 0 f 91 288 10 15 20 25 30 35 40 45 HELIOGRAPHIC INCLINATION, DEG
(B) TITAN III C. Figure 10. - Performance of single 10 kilowatt, 2600 sec- ond specific impulse solar-electric design compared to ALL-Chemical stages. Burner II propellant loading, 1040 kg; electric system total specific mass, 30 kglkW; launch velocity and coast are timing optimized.
9
800 SPECIFIC POWERPLANT MASS - 25 kglkW M U t— ,r-TANKAGE - 3 PERCENT, MID-1970's Y `` THRUSTER EFFICIENCY w 600 E V REFER- \~ ^' ENCE 400 LA_
L SPECIFIC POWER- z PLANT MASS - 35 kg/kW ^^
01 1 1 1 1 1
20 25 30 35 40 45 • V r r HELIOGRAPHY INCLINATION, C14 Figure 11. - Effect of state-of-the-art assumptions on performance. Atlas/Centaur launch vehicle, three- burn trajectories. Reference curve: specific power- plant mass, 30 kg/kW; tankage, 10 percent; 1968 thrus- ter efficiency; thrust control, normal to orbit plane; all free variables optimized.
NASA-Lewis