AIAA 2001-3377 Solar Thermal Propulsion for an Interstellar Probe

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

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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

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

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. The more dissociation of the propellant and thus higher expansion ratio has a slightly greater effect with ISP values maximue Th . m ISP predicter dfo methane in that after an expansion ratio of 70:1, the temperatures of 2400 K, 3000 K, 3300 K and 3500 K increase in ISp is less than two percent. sec0 are86 , 1037 sec 126d , 116an 7c sec6se , respectively observation A . made nb than e tca fro m Methane (CH4) Specific Impulse (Vacuum ) the analysis results is that the nozzle area expansion Relative to Temperature and Nozzle Expansion Ratio Pchamber = 517 kPa ratio doe t havsno largea e IS e affecPth n leveo t t a l ratios greater 50:1. Ther less ei spercen o thatw na t increase in ISp from an expansion ration of 50:1 to 100:1.

Hydrogen (H2) Specific Impulse (Vacuum ) Relative to Temperature and Nozzle Expansion Ratio Pchambea kP 7 51 r=

Figur . Methane 6 Initia a kP 7 l ePressur 51 I St Pa e

Methane (CH4) Specific Impulse (Vacuum) Relative to Temperature and Nozzle Expansion Ratio

Figure 4. Hydrogen ISp at 517 kPa Initial Pressure

Hydrogen (H2) Specific Impulse (Vacuum) Relative to Temperature and Nozzle Expansion Ratio Pchambe 138r= a 0kP

Figur . Methane7 e I St 138p a Initia a 0kP l Pressure

Ammonia produces the lowest ISP levels of the three propellant candidates. Again the analysis 20 30 40 50 60 70 was carried out only at the temperatures of 2400 K, Nozzle Area Expansion Ratio 3000 K, 3300 K and 3500 K. The results of the ISP Figur . Hydrogee5 n I t 138p a Initia a 0kP l analysis (see Figures 8 and 9) for ammonia are S relatively clos thoso et methaner efo . Again characteristic similae othee s ar case o th o rtw rt n si An analysi methanf so propellana s ea t terms of producing higher ISp levels with the lower candidate results in the next best Is? levels of the pressure due to more dissociation. The highest three candidates. This was only evaluated for the predicted levels of ISP for ammonia are temperatures four temperatures of 2400 K, 3000 K, 3300 K and of sec2401 42 , 300 0e 350K d , 330ar 0 K an 0K 0 K

502 sec, 559 sec and 604 sec, respectively. The Table 1 Maximum ISP Values at 100:1 (Minimum expansion ratio ammonieffecte th r fo sclosee aar o rt pressure = 69 kPa, expansion ratio = 100:1) the results of hydrogen analysis. From an expansion Pressure Temperatur) e(K of 50: 100:o 1t 1 ther less ei spercen o thatw na t 2400 3000 3300 3500 increas ISe Pth leveln ei . H2 517 860 1037 1166 1267 69 875 1144 1336 1369T n CH4 517 480 588 667 705 Ammonia (NH3) Specific Impulse (Vacuum ) T TT Relative to Temperature and Nozzle Expansion Ratio 69 485 628 698 705

700 Pcnamber = S17KPa NH3 517 421 502 559 604 69 427 547 634 639m i—— —— ' r 1 500^ ' Pressure = 165 kPa n |400. Pressure = 910 kPa ^r-3000'K tn -^5-3300'K Pressure-221 kPa -*-3500°K In summary baseline th , e propellant hydrogen show mose sth t promis obtaininr efo e gth 0 maximum ISP level resulte Th . s indicat nozzlea e 0 10 0 9 0 8 0 7 0 6 0 5 0 4 0 3 20 Nozzle Area Expansion Ratio expansion ratio of 100:1 also maximizes the ISP, however a ratio of 50:1 is probably adequate and Figure 8. Ammonia ISP at 517 kPa Initial Pressure helps minimiz weighte eth finaA . l design process would optimiz expansioe eth n ratio wit Ie Sphth . As stated earlier, the AV is also a function of Ammonia (NH3) Specific Impulse (Vacuum ) Relative to Temperature and Nozzle Expansion Ratio the mass ratio or mass fraction. The relationship is logarithmi basicalld can y require minimue sth m mass ratio possible. In other words, maximize the throw weight (fuel minimizd masan y ) dr se (taneth k structure, probe, nozzle, effecetc.)e masf Th o .t s

ratio relative to AV for various ISp values can be represented graphically as shown in Figure 10. This show masse thath s ta rati decreases oi leso dt s than 0.5 increase ,th An e i Vdramatics i .

Nozzle Area Expansion Ratio Delta Velocity - AV vs. Mass Ratio - MR

Relativ I eSo Incrementt n P i 0 10 f so Figure 9. Ammonia ISp at 1380 kPa Initial

The result thif so s analysis indicate thers ei an advantage of using lower pressures at higher temperatures to increase the dissociation and thus producing higher ISP values. Further analysis of each propellant at temperatures from 2400 K and higher mads wa determino et maximue eth m ISP usinga minimum (1a pressur0kP psia) 9 pressure 6 f Th eo . e 0 1. 9 0. 8 0. 7 0. 6 0. 5 0. 4 0. 3 0. 2 0. 1 0. 0.0 was adjusted if necessary to reach the temperature. This data is shown in Table 1 with the ISP values at Figure 10. Delta Velocity-AV and Mass Ratio-MR 517 kPa for comparison. The results show that higher value possiblIe f St alwayso par no a t t esbu a lower pressure as the methane results indicate. The The results of this analysis indicate that in 1 ideal pressure for the expected temperature of km-s"orde5 goa1 V f meeA o ro lt e , mastth s ratiof so operation is a parameter in the design that would 0.022, 0.078, 0.148, 0.217, 0.28 0.33d 0an e 5ar benefit fro optimization ma n exercis eacr efo h neede Ir Ssec p0 dfo sec 0 valuesec,0 80 ,60 ,40 f so propellant considered. 1000 sec, 1200 sec and 1400 sec, respectively. A system with a mass ratio close to 0.1 is not uncommon with solid propellant propulsion systems

but is difficult to achieve with the liquid or gas propellane Th t management system (nozzle, heat propellants deep-spaca r Fo . e mission witha exchanger, etc.) woul optimizee db provido dt e eth space-storable bipropellant system, current thrust for the necessary action time. This effort may technology (N2O2+N2H2) yields 317 sec ISP at be necessary in determining the best propellant.

Preliminary estimate massef so maie th nf so THERMAL ANALYSIS components of the system were made to determine an The propellant for the system is stored until

approximate AV and the corresponding VESc values. maneuveV A needee th periheliont r ra d fo thit A .s baseline Th e fue Hestimatesf e uses o l2i th n di . time, heatin requires gi produco dt necessare eth y ISP. Tabl summarizee2 weighte sth included san e sth The solar propulsion concept utilizes the Sun as the mass rati masd oan s fraction. energy source to accomplish this task. For this system panele ,th exteriose useth r dfo r heat shields Tabl . Mase2 s Estimate Summary (see Figur wer) e2 e identifie mose th s td a practica l Probe 50 kg heat exchanger system. In order to minimize Thermal Shield 141kg complexity and weight, the goal was to use only the LH2 Tank 300 kg center heat shield panel. The entire exterior heat g k 9 20 LH2 shiel 5-metea s di r long carbon-carbon cylinder with Nozzle & Plenum 50kg 1.37-metea r radius, providin gsurfaca e are 43.f ao 1 Total 750 kg m2. The heat shield is divided into three sections of 120 degrees, each with surface areas of 14.3 m2 that Dry Mass 541kg fold out to provide protection for the interior Fuel Mass 209kg component systeme th f so . Mass Ratio 0.721 The heat-exchanger concept uses Mass Fraction 0.279 carbon-carbon channels that run parallel to the axis of heae th t shield cylinder shows a , Figurn i e eth 3s a , Base thesn do e estimate5 goa1 V f A o l e sth heat exchanger "core". Inlet and outlet headers, not km-s"1 is not achieved. The AVs using these shown in the figure, would connect the flow to the estimate km-s"8 2. km-s"e 3 1sar 3. , km-s"7 13. , d 1an supply tank and the nozzle. A parametric thermal 1 analysi performes swa determino dt e effece th eth f o t 4.1 km-s" for the maximum predicted ISp values of 877, 1037, 1166 sec and 1267 sec, respectively. various design parameter He 2th propellann so a ( t 1 thermal analysi t includeNH r CHd sno fo aan s i n di These correspond to VEsc values of 8.7 AU-yr" , 9.6 4 1 1 1 this study). The spacing of the channels was one of AU-yr" , 10.1 AU-yr" 10.d an 6 AU-yr" fro R4 msa the variables in the parametric analysis, along with perihelion. A 3 Rs perihelion increases these values by 7.5% assumin massee gth hele sar d constant. numbee th carbon-carboe plief th ro r sfo n material Assuming the volume of the system is the (this effected the "fin efficiencies" of the internal heat only constraint othee ,th r propellant candidates with transfer surfaces). Flow ratincidend ean t heat flux alse ar o include parametrie th n di c study. lower ISp valuee sth coulo t e d du improv V A e eth large decrease in mass ratio. This is due to the The heat exchange modeles rwa d usinga combination of density and ISP characteristics of the finite-difference scheme with 1000 differential sections along the 5-meter length. For each section, propellant. For example, by replacing the LH2 (70 3 3 the local temperatur pressurd ean e were calculated kg-m" ) with NH3 (682 kg-m" ) the mass ratio would decreasee b 0.2althougo dd t 1an maximue hth m based on incident heat flux, inlet gas temperature and inle pressures ga t thir sFo . analysis, surface predicted ISp values of 421 sec, 502 sec, 559 sec and reradiation, conductive wall resistance, and surface 604 sec are lower than those for H2, the resultant AVs km-s"4 6. km-s"e 7 1ar 7. , km-s"3 km-s"6 9. 18. , d 11.an efficiencies were considered. Radiative losses at the This increase in AV values of approximately 230% back wall (interior surface of the heat shield) are 1 based on the (higher) surface temperature of the front increase Ve ESsth C 13.o valuet 3% AU-yr"s52 , 14.5 1 1 1 surface (exterior surface of the heat shield) for a AU-yr- , 15.3 AU-yr" and 15.9 AU-yr" from a 4 Rs perihelion disadvantage Th . thio e t tha s stotai e tth l conservative estimate. Nusselt numbers and friction weight of the system increases to 2577 kg. factors were calculated using standard correlations Whethe hige rth h specific impulse based locan o l Reynolds number (laminar/turbulent). Flow development effect t consideres no wer s ewa d propellan high-densitte (Hth r 2)o y propellant (NHs) is used, the next hurdle is to optimize the system for a due to the long channel length as compared to the minimum dry mass and maximum propellant mass effective channel diameter. while maintaining structural and thermal integrity. The H2 thermal properties required for the analysi basede sar handboon o k valuer sfo

temperature, witK h0 extrapolatio20 o t p su r nfo analysis, discussed below, shows that one-ply of higher temperatures carbon-carboe Th . n thermal carbon-carbon is not feasible and two-ply is conductivity is based on a linear fit for known values marginal. However tese ton , case using one-plf yo W/m-7 6 d 240Kt an a W/m-. K 6 0K 0 28 f 30 o Kt a carbon-carbon is analyzed. A set of mass flow rates Emissivity (e) for the carbon-carbon material is fro mminimua higma o desireht s ratg/ e0 d rat20 f eo assumed to be 0.9 for the initial analysis. The of 2000 g/s is used to bound the possible range. incident heat flu initialls xi y assumea e b o dt Base thesn do e preliminary finding assumptionsd san , 2 conservative 381 W-cm" averagRn 4 st A .a e value the matrix shown in Table 3 was made for the incidenfoe rth t heat flux woul 40e db 1 W-cm"4 t 2a parametric study. Inlet pressures were determined 2 Rs and 713 W-cm" at 3 Rs using a solar luminosity such that the final pressure is at least 350 kPa (50 16 9.33e2f o 5 cal/se R6.96eld f so c an O cm.e Th psi). analysis is also based on the assumption that the incident heat flux is completely absorbed by the heat Table 3. Parametric Analysis Matrix shield. Channel Flow Test Qinc #of Pinlet maximue Th m possible surface temperature, 2 Size Rate Case W/cm plies KPa heae th T tf sshielo , d occurs when reradiation, mm g/s including the back wall radiation loss, equals the 1 381 1 10 200 500 incident radiation maximu.e Thath , tis m temperature 2 381 3 5 200 500 occurs when the net surface heat flux is zero, as 3 381 3 10 200 500 shown in the equation 4 381 3 15 200 500 5 381 3 5 1100 1500 where a is the Stefan-Boltzmann constant, and the 6 381 3 10 1100 500 factor of 2 conservatively accounts for emission from 7 381 3 15 1100 500 bace th k surfac same th t eea temperatur fronte th s e.a 8 381 3 5 2000 1800 fluie Th d temperature cannot exceed this maximum 9 381 3 10 2000 1800 surface temperature incidenn a r Fo . t heat flu 3f x8o 1 10 381 3 15 2000 1800 W-cm"2 and an emissivity of 0.9, the maximum temperature is 2472 K, as compared to 2504 K if the 2 The results of this analysis indicate that the average incident hea W-cm"1 t flu40 f xo useds i . flows above 1100 g/s were too high to allow the fluid However, this would increase to 2891 at 3 Rs, for 2 temperature to reach the maximum surface whic nominae hth l incident hea W-cm"3 t flu71 s xi . temperature. Channel size affected bot pressure hth e surfaca f o e e Us material wit hlowea r emissivits yha asymptotie th dro d pan c temperature pressure Th . e the most potential for increasing the maximum drop is greatest with the 5 mm channel while there is temperature exampler Fo . materiaa , l witn ha the least degradation of asymptotic temperature for longee th emissivit t (a r wavelength3 0. f yo s the same channel size. Result tese th t f caseso thrs2 u associated with reradiation) increases the maximum 4 are shown in Figures 11,12 and 13. The maximum temperatur R3 s, d R 4 380 d an st 329 o a e t an 5K 5 K predicted temperatures achieved with test cases 2, 3, respectively. and 4 are 2471 K, 2448 K, and 2368 K, respectively. parametersf o t Ase valued e th an r sfo The corresponding pressure drops are 90 kPa, 7 kPa, parametric analysis were defined as follows: and 2 kPa. Incident heat flux: 381 W-cm"2 Propellant Temperatur Hean ei t Exchanger plief o # s (thickness): 1 ) (0.3mm Emissivity = 0.9, Flow = 200 g/s, Channel Size = 5 mm, 3 ply, 4 F 2 ) (0.6mm ) (0.3 9mm Channel size: 5 mm 10mm m 15m Mass flow rate: 200 g/s HOOg/s 200s 0g/

2 The incident heat flux of 381 W-cm" is Heat Exchanger Distance (m) selecte evaluatdo t minimue eth m expected heat transfer capability. Results from the structural Figure 11. Parametric Analysi s- Tes t Case2

heat flux. Thi accomplishes si d maneuve witV A ha r Based on these results, the recommended aperihelioa t R3 sf nwhero incidene eth t heat flus xi channel siz 10mms ei thit A .s channel sizee ,th increased to 713 W-cm"2. In order to prevent the propellant temperature is within one percent of the maximum temperature of the carbon-carbon material surface temperature and the pressure drop is minimal. from exceeding emissivit e 350th , hea0e K th tf yo actuae Th l size coul optimizee db would dan d shield needs to be a value of 0.4 or higher. The probably be less than 10 mm. Further analysis was results of analyzing with these parameters are shown performe narroo dt range wth flon ei w rate capability in Figure 15. since the range between the low and middle values originally selecte largeo s s .d wa Thi s analysis shows STRUCTURAL ANALYSIS that flow rates higher than 500 g/s result in excessive The structural analysis was performed to degradatio asymptotie th f no c temperature ar d ean evaluat heae eth t exchanger concep propulsioe th f to n not desirable for the 10 mm channel size. system heae Th t. exchange assumes ri consiso dt f o t heae th t shield center panel onlyheae Th t. shiels di

Propellant Temperatur Hean ei t Exchanger carbon-carbon with channels that run the full length. Emlssivit 0.9y= ,g/s 0 Flo20 , wChanne= s lply 3 R Siz , ,4 1e= 0 mm channee Th l sizes were varied along wit materiae hth l thickness or number of plies for this analysis. A typical configuration of the channeled heat exchanger is show Figurn i e 3.

Propellant Temperature in Heat Exchanger Emissivity = 0.3, Flow = 232 g/s. Channel Size = 10 mm, 3 ply, 4 Rs

Heat Exchanger Distance (m)

Figur Parametri. e12 c Analysi s- Tes t Case3

Propellant Temperatur Hean ei t Exchanger Emisslvlt 0.9y= ,g/s 0 Flo,20 wChanne= s ply3 R , ,4 mm l Siz5 1 e=

Heat Exchanger Distance (m)

Figure 14 Thermal Analysis - Emissivity = 0.3

Propellant Temperature in Heat Exchanger Emissivity = 0.4, Flow = 232 g/s, Channel Size = 10 mm, 3 ply, 3 Rs

Heat Exchanger Distance (m)

Figure 13. Parametric Analysis - Test Case 4

The possibility of increasing the asymptotic temperatur decreasiny eb emissivite gth alss yowa

investigated proposee On . d metho reducinr dfo e gth Heat Exchanger Distance (m) emissivit applo t s yi thiya n (~15u) coatinf go tungsten on the heat exchanger. This concept was Figure 15 Thermal Analysis R 3 s- analyze configuratioe th r dfo n wit hchannels m 10m , 3-ply carbon-carbon materia floa d w2 an lrat 23 f eo The channel sizes that were considered g/s. An emissivity of 0.3 was the lowest value thicknese Th e . th 1f d s0o mm includan m m e5 analyze correspondine th d dan g result showe sar n i materia varies wa l d plie 6 fro o incrementt n si m 1 f so Figure 14. This increase outlee sth t temperaturo et thick m typicae m Th wher.3 y 10. pl l s e eacwa y hpl 3200 additiona n KA . l increas asymptotin ei c carbon-carbon material propertie 3000°t sa F (1922 temperature is achievable with an increase in incident K) were use thin di s analysis pressure Th . e level

assumed for the evaluation is 200 psi (1379 kPa), requirement additionan A . l complicatioe th s ni which make analysie sth s more conservativi ps 5 7 f ei relatively larg heaee sizth tf eexchangeo e th d ran (517 kPa) is the best pressure for the desired ISp. The relatively large mismatc coefficiente th f ho f o s acceptable configurations were determined using thermal expansion for carbon-carbon and tungsten. stress criteria resulte Th .thi f so s analysi showe sar n A second method of achieving the higher in Table 3. temperatures after reducin emissivitye gth througs i , h Table 3. Heat Shield Maximum Stressi ,ps increasing the incident heat flux. This is possible maneuvewitV A ha perihelioa t ra R3 f .n o Thit sno Plief o # s Channel Size s 5 mm m 10m only increases the maximum possible propellant ) (0.y 1pl 3 mm 23,740 91,890 temperature but also decreases the required AV. Any 2 ply (0.6 mm) 4,278 22,450 optimization in this area would need to consider the 3 ply (0.9 mm) 1,996 6,571 mass chang heae th t e n affeces i shiel th it n d o td an 4 ply (1.2 mm) 1,249 4,044 mass ratio since closer maneuvers probably require ) (1.y pl 5mm 5 874 2,779 more shielding. 6 ply (1.8 mm) 658 1,930 ACKNOWLEDGMENT * Allowable stress ~ 14,000 psi Others contributin thio gt s effort includ. eJ Me Adams (mission design) and B. Williams (thermal This analysis shows tha botr fo t h sizef so design concept) ,John e botth t ha s Hopkins channel 1-ple sth y desig unacceptables ni 2-ple Th .y University Applied Physics Laboratory McNut. R . t desig sizm acceptables nm i only5 e th , whilr efo e eth thank . RemeikisS helr sfo p wit manuscripte hth . remaining designs are acceptable for both channel Suppor provides wa t d under Task 7600-039 froe mth sizes. NASA Institut Advancer efo d Concepts (NASA Contract NAS5-98051). CONCLUSIONS 1 overale Th l goaachievo t s AU-yr"li 0 2 ea REFERENCES (94.8 km-s"1) escape velocity fro solae mth r system, 1 1. Jaffe, L. D. and C. V. Ivie, Science aspects of a which requires a ~15 km-s" AV maneuver at mission beyond the planets, Icarus, 39, 486-494, perihelion using the STP system. An initial 1979. assessment suggeste possibilita de thib o e st th f yi 2. Interstellae alt Holzere Th . , E. . T , r Probe: propulsion system could deliver an ISP of 1000+ sec. Scientific objective Frontiea r sfo r missioe th o nt The STP system can deliver an ISp in that range based heliospheric boundar interstellad yan r space, on the assumptions made in this report. NASA Publication, 1990. However, ther mors ei e require achievo dt e 3. Mewaldt, R. A., J. Kangas, S. J. Kerridge, and overale th l goalnexe driveg Th tbi . r afte Ie Sf po th r M. Neugebauer smalA , l interstellar probe th o et the system is the mass ratio. To obtain the mass ratio heliospheric boundary and interstellar space, values discussed, while moleculausinw lo ga r weight ActaAstron., 35, Suppl., 267-276, 1995. propellant, requires the optimization of structural . 4 McNutt . L. R ,, . Gold . RoelofJr.E C . . . ,J R E ,. L , masses and an innovative way to carry and store the Zanetti . ReynoldsL . . E Farquhar, W . . R ,A . D , propellan lone th gr periofo t timf do e before actually Gurnett, and W. S. Kurth, A sole/ad astra: From requiring its use. Since the tank is the heaviest major starse th . Brito ,J t n th Inter.eSu Soc., 50, 463- componen systeme th f o toptimizatios it , e th s ni 474, 1997. highest priority heae Th t. shield (and heat 5. McNutt L.. R ,,. Andrews Jr.B . ,G . McAdamsJ , , exchanger) is the next component in priority order for . RGoldE . . SantoA , . OurslerD , , K. Heeres. M , optimizatio masr nfo s reduction desige Th . n Fraeman, and B. Williams, A realistic interstellar describe thin di s repor vers ti y preliminars ha d yan explorer, STAIF-2000 Proc., 2000. roo sucr mfo h optimizations. . 6 Liewer . MewaldtA . C.. P , . R ,. Ayon R A . d J , an , heae Th t exchanger/heat shield describen di A. Wallace, NASA's interstellar probe mission, the report can serve both purposes but to achieve the STAIF-2000 Proc., 2000. higher ISP values the temperature must be higher than 7. McNutt, R.L., A Realistic Interstellar Explorer, the baselin meane th f achievinf eo s o case e On . g Phase I Final Report, NASA Institute for such a higher temperature is to lower the emissivity Advanced Concepts CP98-01C MA ,1999 y Ma , . from that of carbon-carbon by using a coating of low 8. R. L. McNutt, Jr., G. B. Andrews, J. McAdams, emissive material suc tungstens ha . Processes similar R. GoldE . . SantoA , . OurslerD , . HeeresK , . M , thio t s have been pas e donwoult th tbu n ei d still need Fraeman, and B. Williams, Low-Cost Interstellar to be evaluated and demonstrated for this specific Probe, paper IAA-L-0608, Fourth IAA

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