Results of Evaluation of Solar Thermal Propulsion Gordon Woodcock ... Consultant to Gray Research, Inc. Dave Byers ... Consultant to SAIC ABSTRACT BACKGROUND The solar thermal propulsion evaluation The study was requested by NASA Head- reported here relied on prior research for all information quarters, Code s, to evaluate solar thermal propulsion on solar thermal propulsion technology and for the In-Space Propulsion Technology program performance. Sources included personal contacts with administered by the Marshall Space Flight Center experts in the field in addition to published reports and (MSFC). The study was assigned to SAIC under the papers. Mission performance models were created ISTA contract, which supports In-Space Propulsion at based on this information in order to estimate Marshall. The study was administered by Les performance and mass characteristics of solar thermal Alexander and Bonnie James of the MSFC In-Space propulsion systems. Mission analysis was performed propulsion organization. The study was initiated in late for a set of reference missions to assess the capabilities July 2002, with a completion date of September, 2002. and benefits of solar thermal propulsion in comparison Solar thermal propulsion has been under with alternative in-space propulsion systems such as technology development for about 30 years. The fact chemical and electric propulsion. Mission analysis that hydrogen gas, heated to 2500 - 3000K and included estimation of delta V requirements as well as expanded through a nozzle, could deliver specific payload capabilities for a range of missions. Launch impulse (Isp) in excess of 800 seconds, was well- requirements and costs, and integration into launch known through demonstrations in the nuclear rocket vehicles, were also considered. program. Solar hrnaces are known to reach this The mission set included representative temperature range. It was seen as likely that a solar robotic scientific missions, and potential fiture NASA thermal propulsion system could reach much higher human missions beyond low Earth orbit. Commercial efficiency in converting energy of sunlight to thrust that communications satellite delivery missions were also is possible with solar electric propulsion. The reason is included, because if STP technology were selected for that concentration of sunlight onto a thruster, thereby that application, frequent use is implied and this would heating hydrogen, might have much higher efficiency help amortize costs for technology advancement and than converting sunlight to electricity by solar arrays systems development. A “C3 Topper” mission was and powering an electric thruster with the electricity. defined, calling for a relatively small STP. The app- This higher efficiency, it is argued, would compensate lication is to augment the launch energy (C3) available for the lower Isp, making solar thermal propulsion from launch vehicles with their built-in upper stages potentially competitive with solar electric propulsion Payload masses were obtained from references and capable of much shorter trip times. where available. The communications satellite masses represent the range of payload capabilities for the Delta MISSIONS AND REQulREMENTS IV Medium and/or Atlas launch vehicle family. Results indicated that STP could improve Missions were selected to suit the objectives of the payload capability over current systems, but that this assessment. A list of missions with reasons for advantage cannot be realized except in a few cases selection is given in Table 1. Estimates of performance because of payload fairing volume litations on requirements are given in Table 2. current launch vehicles. It was also found that The current In-Space Propulsion (ISP) acquiring a more capable (existing) launch vehicle, technology program is sponsored by Code S; therefore rather than adding an STP stage, is the most economical the study focused on representative Code S missions. in most cases. Commercial communications satellite delivery missions were included because if STP technology were selected PURPOSE for that application, frequent use is implied and this would help amortize costs for technology advancement The purpose of the evaluation was to assess and systems development. The “C3 Topper” mission is suitability of solar thermal propulsion for in-space a case where the STP is relatively small. It does not propulsion applications, by examining performance and present a problem for payload fairing volume, and its probable cost to customers on a range of representative competition is probably solid propellant, with Isp less missions, compared to current systems. than 300 seconds. The HEDS gateway was selected 1 because it is a mission application for which new in- An existing ISP requirements matrix was space propulsion development is needed regardless of interrogated to obtain destination and payload mass the technology selected, and STP is not at a non- data for the Code S payloads. The GEO comsat masses recurring cost disadvantage relative to other systems. represent the smallest and largest Delta N Medium The "no-hydrogen'' application to an RLV upper stage options. Except for fairing volume considerations, STP is a different competitive environment than ELV upper stages would deliver more payload on the same launchers because the STP does not compete with launch vehicle, but these masses were considered developed cryogenic upper stages. In this case the STP representative. The C3 Topper was examined cannot use hydrogen but is competing with other generically. Scientific payloads for outer planet systems that also cannot. The STP would probably use missions, fiom the ISP requirements matrix, range fiom ammonia as propellant; methane is possible but it is about 300 kg for small, simple payloads such as cryogenic and much more flammable than ammonia. planetary flyby payloads, up to 1500 kg for a Titan The payload masses were obtained from orbiterhander. Even larger payloads may be of interest references where available. The communications at a later time. Examples of greater payload satellite masses represent the range of payload requirements, presently not very quantified, are a large capabilities for the Delta IV Medium launch vehicle Europa lander intended to penetrate Europa's ice to family. The payloads cited in Table 2 are, in all cases, search for the putative ocean below, and a Titan sample the mission payload and do not include apogee return mission. The HEDS L1 Gateway mission propulsion. Communications satellite payloads are payload was obtained from a JSC planning often cited in terms of geosynchronous transfer orbit presentation. The Gateway is a small habitable space (GTO) but the figures here are payload to the mission station. The RLV upper stage mission presumes that orbit, geosynchronous equatorial orbit (GEO). these payloads will utilize the launch capability of an RLV. Smaller payloads may also be of interest. One Table 1: Tabulation of Missions Evaluated such case is launch of an experimental nuclear electric for STP Application propulsion (NEP) stage to LEO with an STP stage designed to transfer the NEP stage to C3=0 that the NGST - Representative small sdence so ..Payload 00. ESLn nuclear propulsion system is not started in Earth orbit. spacel~~omtv ~i-,, * Representative medium sdence ~aykad(to ETSO) This case requires an estimated payload mass to C3=0 TeneStrial planet Finder - - Representative large sdem of 5800 kg. pykad (to ETSO) - MediumGEOComsat * Hiah-demandcommefcial P-Wd MISSION PERFORMANCE AND COST ANALYSIS * LargeGEOComsat - Highdemand commercial payload Barger) C3 Topper for outer planets Smaller STP leads to kss Payload Performance missions volume concern HEDS L1 Gateway - Large payload for EMLl which requires inapace pmpulsion A performance baseline was created for development application to the GEO and Earth escape missions. * RLVUpperStage * #%ohydroge~safety constraint applied. STP with These missions are similar in that both require ammonia may be competitive expanding an initially circular orbit to a highly elliptic orbit, for GEO with apoapse at 42,164 km and for escape or Earth-Moon L1 (EMLI), essentially at infinity, i.e. C3 = 0. For EMLl the C3 is actually about Descrlptlon Deotinatkn Payload Mass Remarks -2 km2/sec2but this is essentially the same from a delta (kg) V viewpoint. For the GEO missions, an apoapse delta NGST EaMunLZ 1400 From ISP Require mantzMatm V about 1800 m/s is required; for Earth escape no Space Mer- Earth Trailing sdar 3900 From ISP Require apoapse maneuver is needed, and for insertion at L1 the (emmetlyWasion Orbit(ETS0) mentsMeMa maneuver is about 650 m/s. High thrust systems can Termstrialplanet Earth-SunLZ 4800 Fmm ISP Require Fi mentsMatm get to L1 or L2 via a powered lunar gravity assist for MediurnGEO GEO 1900 w CaDabilii less apoapse maneuvers (2 required) totaling about 250 m/s canssl I apOgee mot& but STP does not have high enough thrust-to-mass ratio canspt apcgei mot& to perform the gravity assist thrusting maneuver. C3 Topper Outer Planets 300 - 1500 kg Generic capability The simplest way to fly from LEO to these HEDSLIGatway EarthMoonL1 24.000 JSC HEW DRM destinations with STP is continuous thrusting. The RLVUpperStege GEOorC3=0 SMX)kgormore Rationale result is a spiral path away from Earth with substantial ("ydmgen) (NEW issafely G losses. STP does not have high enough Isp to accept these losses; its payload performance would be less than that for conventional chemical propulsion and 2 there would be no