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AIAA 2001-3377 Solar Thermal Propulsion for an Interstellar Probe

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AIAA 2001-3377 Solar Thermal Propulsion for an Interstellar Probe (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. AM A A A01-34143 AIAA 2001-3377 Solar Thermal Propulsion a r nfo Interstellar Probe R. W. Lyman, M. E. Ewing, R. S. Krishnan, D. M. Lester Thiokol Propulsion Brigham City, UT R. McNuttL . Jr , Applied Physics Laboratory Johns Hopkins University Laurel, MD 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference July 8-11,2001 Salt Lake City, Utah For permission to copy or to republish, contact the copyright owner named on the first page. For AlAA-held copyright, write to AIAA Permissions Department, 1801 Alexander Bell Drive, Suite 500, Reston, VA, 20191-4344. (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. SOLAR THERMAL PROPULSION FOR AN INTERSTELLAR PROBE Ronal . LymandW , Mar . EwingkE , Rames . KrishnanhS , Dea . LesternM , Thiokol Propulsion Brigham CityT U , Ralp . McNuthL . tJr Applied Physics Laboratory Johns Hopkins University 11100 Johns Hopkins Road Laurel, MD ABSTRACT readiness and required system improvements will be The conceptual design of an interstellar identified. probe powered by solar thermal propulsion (STP) INTRODUCTION engines will be described. This spacecraft will use Although traveling to the stars is still engineP ST perforo st m solar gravity assist witha beyon reace dth knowf ho n technology proba , e perihelion of 3 to 4 solar radii to achieve a solar capabl penetratinf eo interstellae gth r mediue b y mma system exit velocity (see Figure 1). Solar thermal within that reach.1"6 Usin ghigh-thrusta , high propulsion use sun'e sth s energ heayo t workina t g specific impulse solar thermal propulsion system this fluid wit moleculaw hlo r weight, suc hydrogens ha , probe could potentially achiev asymptotin ea c escape vero t y high temperatures (aroune d Th 300 . 0K) speed of 20 astronomical units (AU) per year. This stored thermal energy is then converted to kinetic mission would utilize a Jupiter flyby and powered energy as the working fluid exits a diverging nozzle, perihelion gravity assist.5'7"9 resulting in a propulsion system with a high specific Solar thermal propulsio innovativn a s ni e impulse (ISp). concept that use Sun'e sth s energ heao yt ta low-molecular weight fluid suc hydrogen.s ha 10"11 thermae Th l energy storeheatee th n di d flui thes di n converte kinetio dt c energ expansioy yb n througha diverging nozzle. This results in a high efficiency propulsion system.12 Spacecraft usin havP geST been proposed for orbital transfer, interplanetary, and other delta velocity missions.13 An interstellar probe using solar thermal propulsio perforo nt mperihelioa n maneuves rha been proposed. probe 5'7"Th 9 e itsel smals fi l (~50kg). Prior to the solar gravity assist, it is surrounded by a much more massive carrier or cocoon. This cocoon protects the probe and consists primarily of the heat systemP shielthermae ST d Th .d an l mode shows li n in Figure 2. This diagram shows a vehicle with a Figure 1. Proposed trajectory for an Interstellar Probe This paper will evaluat feasibilite eth f yo interstellan a r fo usin P grST probe technologe Th . y Thioko200K © * 1AT l Propulsion Corp. Publishe Americae th y db n Institute of Aeronautics and Astronautics, Inc., with permission Figur . Proposee2 d Interstellar Probe wite hth pre-maneuver heat shields deployed 1 American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. deployable exterior heat shield (show redn i ) that rockey b t pioneer Hermann Oberth.14 Measurine gth protects the interstellar probe and propulsion system A km-sec"Vn i asymptotie 1th c escape speed froe mth durin perihelios git n maneuver shielde Th . , madf eo solar syste mapproximatels i y carbon-carbon composite openes i , oriented dan d toward the sun as the spacecraft approaches perihelion. This hinged device will form a 30° yr shado sola4 t w(a r radii thermallo )t y protece tth 74 probe/carrier assembly. wher distance eth rs pi en i fro n centee mth Su f ro The center third of the heat shield is the terms of Sun radii (R$). This equation gives an STP engine. This engine is a channeled heat approximation to the required propulsive maneuver exchanger as shown in Figure 3. During the for a given asymptotic speed. perihelion maneuver, low-pressure hydrogen (< 1380 With this relationship the required AV is kPa) flows through this heat exchanger at 232 g/s for approximately 14.6 km-s"1 for a 20 AU-yr"1 escape 15 minutes. The channels are sized to achieve a high velocity fro R4 msa perihelion. Wit hclosea r heat exchanger effectiveness, resulting in a hydrogen approach R3 , s perihelion requiree ,th reduce s i V dA d exit temperature approachin theoreticae gth l limir fo t to approximately 12.6 km-s"1. However, there are the solar heat source, while minimizing pressure drop additional factors that must be considered for closer of the hydrogen as it passes through the device. approaches, such as thermal protection methods along wit associatee hth d mass increasea s A . baseline, the value of AV =15 km-s"1 was chosen for the propulsion "target enablo "t asymptotin ea c solar 1 system escape of roughly 20 AU-yr" from a 4 Rs perihelion. The relationship between the required AV, systee th mase mth sd I Sratipan o mas(MRe th r s)o fraction (Q is Figure 3. Section of the STP Engine (Channeled 3 1 heat exchanger) where g9s 0i . 8 1x1 0" measureV km-s"A a r fo n d i term km-s"f so mase 1Th .definess i ratio , s dMR ,a masy finae dr rati e initiasr th e o th l(mf th oo r o f)t o l Once the hydrogen exits the heat exchanger, wet mass (mi), and the mass fraction, £, is the ratio of transferres iti d throug plenuha centroidaa mo t l propellane th t mass (mrelationshie mjpo )t Th . p nozzle, where the stored thermal energy is converted shows tha requiree tth directls i V dA y proportionao lt thrusto t shoult I . notee db d tha thin ti s feasibility logarithmia ISPd an , c relationship exists wite hth study the designs of the entrance or exit plenums mass ratio or the mass fraction. To achieve the AV have not been addressed, although a mass allocation maneuve rkm-s" 5 goa1 f o l 1, withou unrealizabln a t e has been made. propellant mass, the ISP must be maximized. An analysi mads swa explor o et e eth PROPULSION SYSTEM possible ISP capabilities of potential propellants. Th systee goath achievo r t ms fo li n ea (Thes propellante ear than si externa n a t l thermal escape velocity from the solar system of 20 AU-yr"1 source heats chemically inert material for expulsion (94.8 km-s" missioe 1).Th 7 n require fassa t transfeo rt insteaP inST providinf do g heat chemically yb Jupiter and flyby, where an unpowered gravity assist "burning "fuel.a ) These include liquid hydrogen reduces perihelion of the transfer orbit. In addition, (LH ), ammonia (NH methand )an , e (CHe )Th . the gravity assist accomplishe plansa e change that 2 3 4 baseline propellant selecte referencr dfo LHs ei since places the trajectory where a perihelion change in 2 potentiae highesth e s th ir ha t fo l t ISP. Each candidate velocity delta , a velocity (AV) maneuver, yielde sth analyzes wa variout da s pressur temperaturd ean e desired solar system escape directionV A e Th . conditions at the nozzle inlet and allowed to expand require periheliot da ndifference isth e betweee nth through a range of nozzle expansion ratios of 20:1 to perihelion velocitie hyperbolie th f so c solar-system- 100:1. Pressures of 517 kPa (76 psia) and 1380 kPa escape and elliptical Jupiter-to-perihelion transfer (200 psia) were evaluated in combination with a set trajectories. temperaturesf o ranging from 1500e K350o Th t . 0K The idea of using a high-ISP, high-thrust feasibilit reachinf yo g these temperature discusses si d maneuve firss wa t identifien Su r e closth 192n do i et 9 American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. later in this report, but it is assumed that at least 2400 ISe P350Th analysi . 0K methanr sfo e (see Figures6 possibls i 350a K d maximu0eK an m allowsa produce) 7 d an s similar finding hydrogee th o st n i margin for carbon-carbon material that melts at term highef so r Is? value lowee th t sa r pressuro t e edu approximately 3800 K. more dissociation. The database used for predicting analysie Th s result Hr 2sfo (see Figures4 ISP, while accounting for dissociation, is finite and sho) 5 highese d wth an t ISp levelthree th f eso this is apparent in the analysis of CH4 at 3500 K. The propellants studied. Result thif so s analyses indicates maximum limite s temperaturi a 343kP o d t 7 051 t ea that at the lower temperatures the pressure does not K. The maximum ISp predicted for temperatures of affect the ISp, while at the upper end of the sec8 2400 58 48 , 300 0343e K d , 330ar 0 K an 0K 0 K temperature range lowee ,th r pressure allowr sfo sec, 667 sec and 703 sec, respectively.
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