From Leo to the Planets Using Waveriders

AlAA, 99-4803 From Leo. to th.e Planets Using Waveriders ‘. A. D. McRonald,J1 E. Randolph., Jet PropulsionLaboratory CaliforniaInstitute of Technology. M.‘J. Lewis Universityof Maryland E. p. Bbnfiglio,J:Longuski .- PurdueUn iversity I’ P.Kolodziej NASA-AmesResearch Center

th International’ Space Planes and Hypersonic Systems arid Technologies Conference. and _ 3rd Weakly Ionized-Gases Workshop l-5. November 1’999 Norfolk, VA

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. FROM LEO TO THE PLANETS USING WAVERIDERS

A. D. McRonald, J. E. Randolph Jet PropulsionLaboratory California Institute of Technology. M. J. Lewis University of Maryland E. P. Bonfiglio, J. Longuski PurdueUniversity P.Kolodziej Ames ResearchCenter

ABSTRACT control surfaces, along with engines and propellant to escapefrom Earth, Other advantagesof this A revolution& interplanetary transportation techniqueover normalinterplanetary delivery methods techniqueknown as Aero-Gravity Assist (AGA) has will be discussed. beenstudied by JPL and others to enablerelatively, short trip times betweenEarth andthe other planets. INTRODUCTION AND BACKGROUND It takes advantageof an advancedhypersonic vehicle known as a waverider that uses its high lift to fly The idea to employ the terrestrialplanets as an energy through the atmospheresof Venus and Mars to sourceusing aero-gravity-assist(AGA) maneuversto provide exceptionally large velocity changesusing significantly increasethe velocity of interplanetary gravity-assist maneuvers. The concept has been spacecraft was first discussed for the Stat-probe understudy in a joint programbetween JPL and the mission to the sun] in 1982. Some preliminary University of Maryland for almost a decade. More -analyses were then completed by Longuski2 that recentlyPurdue University andNASA Ames Research suggestedthe necessity of a very high performance Centerhave also become partners. The waverider aerodynamicvehicle having a lift/drag (L/D) ratio of concepthas been proposed as an upper stagevehicle almost 10. This high performanceat the extreme compatible with the ‘LockheedMartin Venture Star Mach numbersnecessary for AGA maneuversseemed SST0 vehicle. This integratedvehicle conceptcould unattainableat that time and the concept was not be used to launch spacecraft on interplanetary pursued. In the mid 198Qswe becameaware of an missions. The paper will discuss the mission aerodynamicvehicle concept introduced by Nonweiler possibilities enabledby a waveridervehicle as well as in the 1950~~which had beenunder study for some the necessarydevelopment program. time at the University of Maryland.4 The first public presentationofsfhe AGA conceptwas made in 1989 Flight durationsto various destinationsin the solar by McRonald and soon thereafterthe connection systemcan be reducedby largefactors (e.g.,2-5 times between the AGA conce t and waveriders was shorterduration). The paper will presentresults of publishedby Lewis, et al7H) ’ from the University of recent studies of interplanetary and atmospheric Maryland. This began a lengthy collaboration trajectories to many bodies, with navigation errors betweenJPL andthe University of Maryland, which andmake-up, of velocity loss at eachAGA maneuver. is ongoing today, and representsthe nucleus of the Vehicle design includes possible ablation of heat teamthat continuesto pursuethe AGA idea for future shield material, andpossible location andnature of interplanetarymissions.

A.D. McRonald is a SeniorMember of the Technical Staff, J.E. Randolphis the Study Managerfor the Solar Probe,both at JPL, California Institute of Technology. M.J. Lewis is a Professorin the Department of AerospaceEngineering, University of Maryland,E. P. Bonfiglio is now a Member of the Technical Staff at JPL, and was. previously a graduatestudent of ProfessorJ.Longuski, who is at the School d Aeronauticsand Astronautics, Purdue University, and P. Kolodziej is a Member of the Technical Staff at NASA Ames ResearchCenter. The researchdescribed in;this paperwas carried out by the .Jet Propulsion Laboratory,California Institute of Technology, under a contractwith the National Aeronauticsand Space Administration. The paperis to be presentedat the AIAA 9th InternationalSpace Planes and Hypersonic Systemsand Technologies Conference at Norfolk, VA on November 1-5, 1999. Copyright 1990by the American Institute of Aeronauticsand Astronautics,Inc. The U.S. Governmenthas ’ a royalty-freelicense to exerciseall rights underthe copyright claimed herein for Governmentalpurposes.. All otherrights arereserved by the copyright owner. 1 (c)1999 Ayerican Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. ..

More recent publications 9*10”’ document detailed Figurei 1 illustrates a planetary waverider:launched to . investigations of actual AGA trajectories.to various HE0 on a Delta II 7925 or carriedto ’ LEO on the planets. Also, new analyses of specific waverider Venture Star. In both cases the waverider would designsappropriate to CO2 atmospheres(e.g., Venus. havea LH-LOX linear aerospikeengine to producea and Mars) have led to more optimized waverider C3,of ,over50 km2ls2, entailing a delta-V of over 5.2 shapes.A model waveriderbased on thesewas tested km/s from LEO and over 2.0 km/s beyondthe Delta in 1993and the results ‘s were very well predictedby from HEO. the analytical codes. WAVERIDER DESIGN Beginning in 1997, NASA first directly fundedthis research.This led to the formation of an AGA team Waveridersare supersonicshapes in which the bow consisting of Mark Lewis at the University of shockwaveis directly attachedto the leading edge. Maryland, JamesLonguski andEugene Bonfiglio of This meansthat all of the flow that passesthrough Purdue University; and Angus McRonald, James the shockwave on the lower lifting part of the Randolph,and Robert Miyake at JPL. One of the waverideris containedbelow the .waverider.This has more significant productsof this team effort was the the benefit of producing a generally high value ‘of developmentof a more’ detailedparametric. “ mission availablelift/drag ratio (L/D) at high Mach numbers space”for AGA trajectorieswithin the solar system with lhigh lift, and reducing cross flow and non- using a new computer code. A program called uniformities on the compressionsurface. STOUR14had been developed to parametricallystudy interplanetary . trajectories with gravity assist Wavetiderswere first definedby Nonweiler.’ They maneuversat intermediateplanets. This program was are generatedby starting with a known flow with a modified by the team membersat Purdue University given shockwave; a stream surface parallel to the to incorporateAGA maneuversat each intermediate direction of flow under the wedge is selected to planet. The resulting new code (STOUR-AGA) representthe- lower surface of the waverider. The’ formed the basis for a. Master’s thesis at Purdue by intersection of that lower surface and the original Bonfiglio.‘*. This new program has been used to shockwavedefines the leading edgewith an attached explore in detail the parametrictrajectory spacethat shockwave.This whole process works becausethe exists for the AGA conceptas a function of launch flowfield is mathematically hyperbolic, so that the energy, waveriderL/D ratio, and which’ intermediate carved-outsection which forms the waveridersurface, planets will be chosen for a near optimum AGA representingperhaps a small portion of the original trajectory to a destination. flowfield, still retains the propertiesof that flowfield even:though the generatingbody has been ignored Another teaming relationship has recently developed oncethe waverideris defined. .asa result of a new NASA initiative that has formed an Aeroassist Team composedof members from the Other generatingbodies can also be used as the NASA centers who are developing a future starting point of the waverider flowfield design technologyplan for aeroassisttechnologies, including process.Conically-derived waveriders have been used the AGA technology. As a result of this initiative, a extensively because they tend towards higher new group from NASA Ames Research Center volumetric efficiency than the wedge-derivedforms. (ARC) has beenadded to our AGA team. Members Combinations of cones and wedgeshave also been of this group bring to the team expertisein high exploredfor creatingthe generatingflowfield. ’ For a temperaturethermal protection systems (TPS) and given flight Mach number, both the wedgeand cone- I materials,as well as an experiencein flight testing of shapedforms have only one degreeof fresdom: the aerodynamicentry systems. The new materials that oblique surface-angle. Burnett and Lewis showed are underdevelopment as part of the SHARP program that conically-derivedshapes can also be optimized at ARC promise to perform well in the extremeTPS with volumetric considerations,to produce vehicle environmentof AGA maneuvers. forms that strike an acceptablecompromise between aerodynamicsand packaging, and with realistic, This paper will include the latest results from the roundedleading edges. In fact, nearly any shapethat waveriderconfiguration designs from Maryland, the has associatedwith it a shockwave and supersonic parametric trajectory studies,from Purdue, the TPS downstreamflow can be usedas the initial generating materials research at NASA ARC, and the ‘body for a waverider. In turn, each generating atmospheric flight simulations that have been flowfield containsan infinite numberof streams completedat JPL.

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&aces which can be selected to form the final Hence,what is truly desiredis a generatingflow field waverider,so there is greatflexibility in the process, that can.generatea waveridershape with the positive which leadseasily to the applicationof optimization. attributes of the waveridersderived from the two ,’ ._ g&&rating flow fields. One option is to use a hybrid Considerationsother than simple aerodynamicsand geometry for the generatingshape. For instance, a volume can be included in waverideroptimization. hybrid cone-wedge shape has been used to For instanceTarpley andLewis.i8 showedthe effects successfullygenerate a waveriderthat has a uniform of steady,state flight and static margin constraintson forebodyflow, like a wedge-derivedwaverider, but the optimized design for an engine-integratedwedge- with a good volumetric efficiency, like a conically- derivedwaverider. The performanceof the optimized derived shape.16A flight test of a shapeof this type vehicle, which included the steady state flight is plannedas part of the NASA Ames’ SHARP constraints,’ is significantly greater than the program.The hybrid shapeintroduces extra flexibility performanceof the vehicle optimized with no steady into the waverideroptimization processby allowing state flight constraintsassuming that control surface the optimizer to expandor contract the dimension of deflectionsmust be includedfor trim.. the wedgesegment relative to the cone, and therefore make the flow field more wedge-like or cone-like. Ideally, the design of an AGA vehicle, should be However, since this method requires a three optimized for the entire trajectory, from atmospheric dimensional Euler calculation for each generating entry to exit back to space. For a multiple-pass body, it is not particularly well suited for preliminary AGA, which may involve flight at several different vehicle designstudy, which may requireexploration sustained Mach numbers in different planetary of numerousdifferent shapes. Most recently, work atmospheres,it is not even obvious which is the best has concentratedon an- even more ,promising design point to select. Although -software and the techniquewhich eliminates the need to choose a hardwareare available to accomplishthis task, from a generatingbody andpermits direct specificationof the practical stand point. it remains a challenge to desiredshock wave instead. This is- the so-called implement. osculating (Latin for “kissing”) cones waverider method developed originally by Sobieczky and In this study, an AGA vehicle had been designed coupledto an optimizer by Takashimaand Lewis.” .. using a so-called“ osculating cone” waverider,with a flowfield constructedfrom numerousslices of conical In the method of oscmating cones, the generating flow. The waveriderhas beenoptimized to fit within flow is definedby a design Mach number, a bow ’ a constrainedvolume, correspondingto a launch shock angle,and a shock wave shapeat the exit plane vehicle enclosure. The method of generating an of the waverider;hence, the methoddoes not requirea osculatingcone waverider is describedbelow. generatingbody to be defined. The flow field behind the non-axisymmetric shock is determined by Given that hypersonic waveriders are generally assuming“ locally conical” flow in the normal planes designedusing an inverse processin which a flow along the shock curve. The “locally conical” flow is field is first selectedaround a chosengenerating body, definedby,an osculating slice of flowfield. The shock an intriguing ‘question .is: what is the “best” angleas well as the Mach number which define the generating shape to use in forming ,that initial ilocally conical” flow are kept constant in’ each generating flow field? Since the waverider is osculatingplane to ensurea smoothcontinuous lower generatedonly from a portion of the flow. field, the surfaceon the generatedwaverider shape: The vertex characteristicof the generatedwaverider shape may of the conical flow field in eachplane is determined not necessarily reflect the characteristic of the by the local radius of curvatureand the shock angle. s generating shape. In contrast, waverider shapes The shock curve is chosenso that the change in the generatedfrom ax&symmetric flow fields have better radiusof curvatureis continuousalong the curve, and aerodynamic performance with greater volumetric. a seriesof planesis usedalong the shock curve in the efficiency; however, the flow coming off the exit plane to fully define the flow field behind the forebodyis no longer tmiform, andfor a given inlet bow shock. width the mass’capturearea is smaller. Furthermore, an axi-symmetric forebody/inletcombination requires Note that in the limit of infinite radius of curvature, greaterturning anglesand contraction ratio for a given the conical flow degeneratesinto a .wedgeflow, so pressurerise, which can increasethe cross-sectional that by prescribing a flat shock curve, a wedgecaret- . areaof the vehicle. wing waveridercan he generatedwith the osculating

cone method. Similarly, a conical waveridercan be Engineering, is used to optimized the design for c generatedby simply prescribing a shock curve with maximum cruiserange performance. constantradius of curvature,i.e:, a circular arc. Thus, by choosing a shock curve which has an infinite: The gradientsused in the algorithm are calculated radiusof curvaturealong the centerregion and a finite using finite differencesby the optimization code. radius of curvaturealong the outboardregion, the two. Even for a highly interdisciplinary design,the number positive attributesof the two generatingflow fields, a of variablesrequired to definea waveriderare limited wedgeflow anda conical flow, can be combinedin a and the objective function is readily calculated by single waverider.Since the flow field is assumedto analytical means,so the cost of calculating the finite be locally axisymmetric, the methodcan be inaccurate differencegradients is not severe. when large spanwisepressure gradients are present; however, such a flow would have a corresponding Application .of optimized waverider forms was shape with large, surface’curvature, which would describedby Lewis andMcRonald. ‘* Figure 1 shows likely not be suitable for practical applications. a representative waverider vehicle design. To ’ Moreover,the integratedaerodynamic forces calculated accuratelypredict the surfaceflow properties,a hybrid by the method matched well with the values methodof tangent-wedgeand tangent-cone methods is calculated numerically. More importantly, the used. For the rest of the airframe, the surface flex.ibility provided by the techniquehas alreadybeen propertiesare calculatedusing the shock-expansion shown to produceshapes with superior aerodynamic method. The viscous forces on all the surfaces ate and volumetric performancecompared to the simpler calculatedusing a referencetemperature method. The wedge- and cone-derivedforms, and the greater flow ‘is assumedto be entirely turbulent with constant flexibility also enables the designer to place a wall temperature. waverider aerodynamic shell around an existing vehicle with specifiedgeometry and volume. The waverider in Figure 2(a) was optimized for maximum L/D, and to fit within a Venture-star-class Once the generatingflow field is established, an payloadbay. Maximum L/D is 8, and volume is 17 osculating cone waverider shape is determinedby m3. Figure 2(a) shows this shape and Fig. 2(b) selecting a leadingedge ate the exit plane, subject to showsits aerodynamicperformance (CL, Cd, L/D) as certain limits of permissible geometry. When the a function of the angle of attack alpha, for turbulent vertex of the osculating cone and the leading edge- boundarylayer conditions. Figure 2(c) and Figure point is determined,the lower surfaceof the waverider 2(d) show the sameparameters for a laminar boundary can be constructedby tracing the streamline along a layer: Note that below the design attitude, L/D falls known conical flow, ana1ogous.mthe generationof a off rapidly, but at large angles of attack (AOA), L/D conical waverider.If the shock curve is flat, i.e., the is relatively insensitive. radius of curvature is infinite, the slope of the streamlinewill be a constantvalue which is equalto AUTOMATED SEARCHES FOR the wedgeangle which producesthe given bow shock AGA TRAJECTORIES anglefor the given designMach number. ‘Using a constantL/D assumption,an AGA algorithm One of the most important characteristics of the was developedthat accountsfor drag andwas suitable waveriderdesign is that it is an inverseprocess, where for installation into a program known as STOUR a desiredflow. field is first selectedand then ,a shape (Satellite Tour Des$n Program). STOUR was which producesthe generatedflow field is determined. originally developed at JPL where it was used For any given generating flow field, an infinite interactively to designthe Galileo Orbiter Tour. The number of waveriderscan be selectedby varying the programwas upgradedat Purdue’toperform automated shock-intersection curve. For the osculating cone designfor a’variety of gravity-assist missions.*’ A waveriderdesign,. the shock wave shapecan also be. detaileddescription of the AGA algorithm that was varied. Thus, an optimizer can be usedto select the used in the latest version of the program kwon as “best” shapeamong many, where “best” .is definedby STOUR-AGA can be found in Bonfiglio, Lon*quski, some objective function that can relate either to the and‘ Vinh.21’22 Sims, Longuski, and Pate1. and pure aerodynamicform or an integratedperformance Bohfiglio2* used an automated search method to parameter.‘ In this study, the sequential quadratic determineAGA trajectories but assumedan infinite programming method a s implemented by Design ‘LID ratio andlater provided analytic approximations Optimization Tools (DOT) available through VMA of drageffects. , (c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. .

Thereare many potential missions that could benefit conclude from these comparisons that AGA can from AGA trajectories.A Neptune orbiter might be provide an advantageoustrade-off -between time of very interesting, especially if AGA is able to reduce flight andtotal launchenergy. the launch energy and time of flight reqiii&d.: ;.a.. .! ./ ‘Zj, Currently, a numberof missions arebeing planned to Now that we have shown AGA to outperform pure Mars. Cheaperfree-return trajectories that can obtain gravity-assist missions in terms of launch Vm and atmosphericsamples would provide valuablescience. time of flight, it is useful to determinethe best AGA AGA sample-returnmissions from the atmosphereof trajectoriesfor various missions. (A more detailed the gas giants (such as Saturn) could be used to search is provided in Bonfiglio, Longuski, and determinethe age of these distant planets. The Vinh).21*22When designing a mission. the.definition missions that were investigateduse Venus and Mars of “ebesti” can sometimesbe ambiguous. The best for AGA. Randolph and McRonald” .found the AGA trajectorieshere have,been detemrinedon the combination of Venus and Mars for AGA to be basis of launch V,, time of flight, and to a lesser I extremely beneficial, becauseof the lower launch extent arrival V. Table 3 gives a list of the most energy to the first planet (Venus), and the lower promising trajectories for the missions mentioned periapsespeed of the secondplanet (Mars). above. We see from this. table that trajectories to Pluto and Neptuneexist with times of flight less than One mission of greatinterest for the past few’ decades 5.5 years,with an L/D of 10. Additionally, the Mars was a Pluto flyby. Pluto is the only planet that has free-returntrajectories give very promising results. not been visited by spacecraft.-For this reason, a’ mission to Pluto is very exciting and scientifically Patel, Longuski, and Sims25*‘6did extensiveresearch important. The current referencemission to Pluto on Mars free-returntrajectories using. a launch date (now termedPluto-Kuiper Express) uses an 8 year’ range of l/1/1995 to l/1/2020. They showed that trajectory with a large launch vehicle (C3 about 150 trajectorieswith a 2-yeartime of flight and a launch km2/s2) and a gravity assistfrom Jupiter.24Figure 3 V_ of at least 6.0 km/s exist approximately every shows AGA results for a trajectory search with two years. They found trajectories with very fast similar characteristicsto the referencemission. Table times of flight (about 1.5 years), but with a higher 1 gives a legendfor understandingthe STOUR-AGA launch V, of 7.0 km/s, and high arrival V, results. The assertion for our .typical AGA (between8 km/s and 10 km/s). Thesetrajectories do trajectories is that both Venus and Mars AGA not occur often, existing only in 2000, 2002, 2015; maneuvers are necessary. Thus the “path” of and 2017 for their 25year’ search.If a very low L/D preferencewould be “3 2 4”, followed by the final ratio of 3 is used for the free-returnpath, shorter _ body as shown in Figure 3 (i.e.;’ path 3249 goes to times of flight exist with even lower iaunch V; Pluto). A reasonable L/D ratio for the waverider values. An independentre-creation of the Patel, developmentat this time would havea valueless than Longuski, and Sims fast trajectories (with STOUR- 10, with a typical value of 7 at both Venus and Mars, AGA), shows the arrival dates,are identical to the as shown in Figure 3. With this L/D the parametric arrival datesfor the Mars AGA trajectorieswith times variable of Figure 3 is the launch hyperbolic excess of flight of 1.0 year. This comparisonsupports the velocity (Vinf) which should be minimized to enable notion that the best AGA trajectories are the small (inexpensive) launch vehicles to be used. -trajectoriesfound by Patel, Longuski, and Sims, but Table 2 provides ‘example data from two AGA are improved by AGA (reducinglaunch V, as much trajector$ for comparisonagainst the 2004 reference as 4.5 km/s). The arrival V, for the Mars AGA mission. In the first example,we seethat if we use trajectoriesis still high, but is alleviated by using the the exact samelaunch V_ for an AGA trajectory (12 VM path for the free return. Using the VM path, a km/s), then the time of flight is reducedby 3 years time of flight of 113years can be obtainedwith a (comparedto the g-yeartime of flight of the baseline launchV, of 4.0 -km/s with much lower-arrival V,s mission). But, of course,this implies a very large (e.g. 3 km/s vs 9 km/s).comparedto the best time of launch .vehicle, which is not the goal of AGA. flight casefor a pure gravity assist (using ME). However,if we match the g-yeartime of flight with an AGA, we can reducethe requiredlaunch V, by Finally, the Saturn free-return trajectories provide ~- about 4.5 km/s, which means that a smaller and short flight times and low launch V, values. The cheaper‘ class of launch vehiclescould be used. We arrival V, values at Earth are fairly high, but the ability of the waveriderto perform aerocaphtreupon return to the Earth alleviatesthis problem.

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Legendfor STOUR-AGA Launch-DatePlots

PATH Planets encountered,including launch and destinationbodies, e.g. PATH: 3 5 9 usesa Jupiter gravity assistfrom Earth to Pluto vinf Launch V-s. The numerals 1, 2, 3,4 ,...on the plot representthe 1sr,2nd,3rd,4rh,. . . V-s in the list. If an AGA maneuver is used during the trajectory then, A,B,C,D ,... is used instead of 1,2,3,4,... j . E.g. in Figure 3 the ‘D’ on the plot ‘denotesa V-s of.12 km/s. For an STOUR-AGA plot, it is possible to have both SAGAand pure gravity assistpoints on eachplot. Lift/Drag Lift/Drag ratio used at each AGA planet.’ If Lift/Drag=0 then there was no AGA performed at that planet.”E.g. in Figure 3 the Path=3 2 4 9 and .theLift/Drag=O.O .O7.0 0.0 meansthere is no AGA maneuverat the first and last planet in the path, but AGAs with Lift/Drag=7 are possible-atthe 2nd and 3rd planet in the path.- SearchEvent Event Path for which data are plotted. E.g. in Figure 3, TOF to Plluto is plotted since encounterwith Pluto’is the 4” event in the PATH. ALTMIN Minimum flyby altitude permitted in the STOUR-AGA run. SearchMin Trajectorieswith flyby altitudesbelow this value are not included in the plot Launch Dates Launch-daterange (YYMMD D) used in the STROUR-AGA run. E.G. 05/01/01 .Searched meansJanuary 1,2005. The launch-date,Increment is also grven, for example “by 15 days. TFMAX Maximum allowable time of flight plotted.

Table 2 ! Pluto Trajectory Comparisons ReferenceMission ComparableAGA Trajectories Time of Launch V-s Time of. Lift/Drag , Launch V,s Flight (km/s) -- Flight : Ratio &-W (years) (years) 8.0 * 12.0 5.0 12.0 8.0 ; -_ 7 7.45

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4 Table 3.

AGA Trajectories for Various Missions

Path Launch Date TOF Launch V, Arrival V, Lift/Drag (MM/DD/YY) ‘(years) (k&s) Mb) Ratio Pluto VM 10/31/13 5.2 ,9.0 30.5 -10,

VM 10/16/13 7.26.7 . 9.08.0 21.122.8 .: Neptune VM 3127106 5.4 7.0 28.2’ .’ 70 VM 3127106 6.2 7.0 24.1 7 VM 3127106 7.7 7.0 18.2 5 Mars Free-Return

VMM ” 8/l3128114 8102 1.32.4‘ 4.03.5 -3.14.5 : Saturn Free-Return VMS 512l/O7 12.5 10 VMS 512107 z-i :-: ._ 19.7 7 (. VMS 5/21/07 6:4 : 6:0 22.2 5 a The only AGA is at Mars in this trajectory.

ATMOSPHERIC FLIGHT PARAMETERS (1) Figure 4 (a and b) illustrates the rotation of the Vinf vector that changesthe heliocentric vector that is inherentin AGA. This section gives some example parametersfor AGA flight in the atmospheresof -e v,‘. AV, I V, = - (2) Venus .and Mars. Since the heliocentric velocity L l+y gain from AGA increaseswith the size of the Vinf r,: (11 vector it is appropriateto considersubstantial Earth 5.” Vinf values, such as 7 km/s (C3 = 49 km2+2). Figure 4(c) illustrates typical AGA trajectory whereV, is the circular velocity, p is the outcomesusing Mars as the intermediateplanet. A atmosphericdensity V, is the velocity at infinity. retrogradeflyby allows a return to the inner solar system (e.g. earth), whereas,~adirect flyby headsfor Figure 6 shows (a) the value of BL = m/CdA, where the outersolar system. Typical AGA parametersfor CL is the aerodynamiclift coefficient, m is the AGA at Venus andMars are shown in Fig. 5 (from mass, assumedfor now to be 1000 kg, A is the ’ Ref. 9)! (a) velocity, (b) bending angle, ,planform area,taken as 114 m2, and Cd is the drag (c)atmosphericflight time, and (d)travel time, as a coefficient, and (b) CL/Cd (L/D) as a function of function of Earth launch velocity, VinfE. The line at angleof attack, alpha. One can see that the’ AGA VinfE = 7 km/s on Fig. 5 (a) indicatesthat VinfV flight should stay within the rangeabout 1 to 5 deg would be about 15 km/s, that Vp (the periapsisspeed) of alphato maintain nearmaximum L/D. at Venus would be about 18 km/s, while at Mars the corresponding values would be about 25 km/s, Figure 7 shows the equilibrium, atmosphericdensity assuming no dragloss at Venus. It will be shown in AGA, on the left handscale, as a function of BL = below that the drag loss at Venus is about 3.5 km/s, m/CLA on the top scale, shown by the solid lines for and L/D = 5, and one should adjust the VinfE Venus andMars. The assumedvalues of m and A accordinglyto find V at Mars. are m = 1000 kg, and A = 114 m2, so that CL. follows from a given value of BL. The scales of FLIGHT ALTITUDE AOA alpha at the bottom are the values from Fig. 2(b) (turbulent)and Fig. 2(d) (laminar) corresponding The flight altitude for AGA and the velocity loss due to BL acrossthe top. Within the figure vertical lines to dragcan be evaluatedfrom two equations: from the AOA scalesmap out a region of density for

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level flight at,Venus andMars with level speedsof LAUNCH CONFIGURATIONS 12, 14, 16 and 18 km/s. The horizontal lines going to the density scale span.the region for AOA 1 to 5, For launch and flight computationsa mass of 1000 deg (lift downwards)for the laminar boundarylayer kg has been taken as nominal. For a vehicle of case. To the left of Fig. 7 there are scalesshowing length 60 ft, (18.3 m), width 15 ft (4.57 m) and the altitude at Venus andMars correspondingto the height 10 ft (3.05 m) (fits in the STS cargo bay) the density (Ref.*‘, *‘): For example, the altitude range surfacearea is about 114m2 for eachof two surfaces. for Venuscorresponding to AOA 1 to 5 deg and V =, Taking a carbon-carbonmaterial (specific gravity 2.4) 12 to 18 km is about 107 to 102km, and for Mars it of thickness I mm would give a mass of 550 kg. is about59 to 51 km. The bottom scale on Fig. 7 is The launchto LEO could be on the STS or a vehicle stagnationpoint heatingrate, and the diagonaldashed like the Venture Star, as shown in’ Fig. 1 (b). To lines marked“ heating” show the convective heating achievea C3 of 49 km2fs2 would requirea delta V of p rate, qref, 2g*30~s’) to a cylinder of radius 0.3 in. It about5.3 km/s,‘requiring an initial mass (with LOX- is well known that qref varies as the inverse square LH propellantat Isp = 44.0set) of 3380 kg to deliver root of the noseradius (Rn), and that the heating to a 1000 kg. Thereforethe tanks and nozzle for the cylinder is less by squareroot of 2 than for a sphere assumedlinear aerospiketype engine would have a of the sameradius. The figure has a vertical line at mass’of perhaps240 kg. An alternativelaunch with 370 W/cm2, the flux that a body at 3000 K could a Delta II (see Fig. l(a)) from HE0 (C3 = -2 radiatewith emissivity 0.8 (a typical infraredvalue). km2/s2) would requireabout 2.0 km/s, giving 1630 The point shown on the line as a circle is a radiative kg initial mass to deliver lOOOkg,and perhaps160 kg. flux calculatedfor Rn = 0.1 m, V = 18 km/s, andthe of tanks and nozzle. To give a worthwhile payload density shown. At high speedthe radiativeheating and the necessaryavionics for the AGA flights, is expectedto exceedthe convective. At a flux below thermal protection,etc., would evidently requiremore 370 W/cm2 a high-temperatureradiator (described than;1000 kg, but this value suffices to illustrate the below) would not ablate, and abovethis ‘temperature designprocess. _ therewould be ablation. If ablation is expectedone might make.the leadin edgeof carbon-phenolic,used CONTROL on the Galileo probe,$2’ which is a charring ablator capableof acceptingmuch higher heatingrates. In To fly at high speedin a corridor of a few km in the Galileo case the predictedpeak radiative flux to altitude calls for rapid determinationof the ambient the probe nose was -over 40 kWlcm2 for a few a&o!lift accelerationforce, and quick implementation seconds, and .a longer convective heating pulse of a’pitch correctionmaneuver. The obvious method peaking at over 17 kW/cm2. The total ‘time-integral of control is a control surfaceat the rear (where the was about 700 kJlcm2, causing ablation of an linear aerospike engine takes some of the space), estimated2 inchesof material,and to reach this value which would changethe effective AOA a degreeor so for Venus AGA, of durationabout 500 set, the heat by moving only a small amount (e.g., 10 % of the flux would be 1.4 kW/cm2. Also, the thickness of chordmoving through5 deg,probably using ram-air- leadingedge material could ,be severalinches, so that derivedpower. We notethat the aeroheating on the both the local flux and the time integral from the wetted (top surfacein AGA inverted flight) will be.a Galileo probe were considerably above the AGA few,percent of the stagnation level, and the heating requirements.Presumably a different waveridershape rate,on the other surfacewill be even less. A rudder would be called for with a charring ablator. Also, or wing-tip surfacesmay be appropriate for yaw one must considerthe thickness of high-temperature control, Control will be facilitated if the center of, thermal insulator necessaryto keepthe interior of the gravity (CG) can be kept well forward of the center vehicle within boundsduring a flight of 500 set at of pressure(CP), which is expectedto be at the Venus,.and200 set at Mars. centroid of area. Having enginenozzles along the rear brings the cg back, but the substantial leading From Eq. (2) for Venus and a Vinf at .entry,of 15 edgethermal protection mass would bring it forward. km/s, we can evaluate,delta V to be about 3.5 km/s, i.e., the Vinf at exit wiil be about 11.5 km/s, and.the ENTRY andEXIT Vinf at Mars will be reducedin accordancewith the data of Fig.5 Some entry dataare given -in Table 4, showing entry angles for various constant L/D entry paths giving

8 (c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or authpr(s)’ sponsoring organization. e

Y level flight close to the equilibrium conditions. It the pass that will be targetedto the, next planet. should be noted that entry andexit occupy a certain Figure 8 illustrates how travel time from Venus to angularrange, and that a certain entry angle error:br,. ~Maiii;Svaries with V and exit angle (a), and how uncertaintywill be presentbut significant corrections angular rangefrom Venus to Mars varies.with V and arepossible by maneuveringin the atmosphere. The exit angle(b). From data of this type an adaptive most likely entry configuration is inverted with algorithm would be on-boardfor AGA at Venus,-to ‘downwardlift at about a mid value, but one must determinewhere to exit from the measureddrag, to considerthat there-ismore lift available if the vehicle reachMars with the shortesttravel time and the least goes initially below the equilibrium altitude. The delta V correction, given the ambient losses in the exit point should be adaptablein terms of the VenusAGA. \ measureddrag accelerationand its time integral, to come out with the highest V at an angular rangeof

Table 4. Some Venus entry cases close to desired level flight (102 to 107 km altitude); constant m/CdA, L/D assumed from entry to level-off

Level-off Entry m/CdA, Entry Angular . Altittude, Speed, km/s kg/m2 angle, deg range from km AtZ= 125 entry, deg : km 102.3 ‘12 500 0 4.0 6.2 106.9 12 500 0 3.5 102.0 12 500 +3 105.7 200 +3 t-i 5.1. 106.8 :;. 200 +5 4:o 4.7 102.2 . 14 500 0 4.2 10710 14 500 0 3.7 105.9 14 500 +3 4.0 4.7 108.5 14 500 -3 3.5 ,- 107.8 14 500 -5 3.5 102.0 16 500 _ 4.5 107.2 16.’ 500 i 3.7 103.4 16 500 +3 4.5 103.8 16 500 -3 4.0 4.0 102.3 18 00 0 4.5 107.0 18 .500 0 4.0 t 4.7 ,104.l 18 500 +3 ., 4.5 105.5 18 200 +3 4.5 106.5 18 200 +5 4.5

TECHNOLOGY DEVELOPMENT NEEDS by numericalsimulation of the coupled aerothermal heating / thermal responsebehavior of the UHTC An enabling technology for aero-gravity assist leadingedge to define the steady-state“ non-ablating vehicles is the ultra-high temperature ceramic performance”constraint on the altitude-velocity flight (UHTC) materials under development at NASA envelope.“” An .APC ‘for the 0.141 inch radius Ames. UHTCs have a unique combination of UHTC nosetip tested in the SHARP-B] flight mechanical, thermal, and chemical properties that demonstrationis shown in Fig. 9 alon? with. a enablethe fabrication of very small radius or sharp typical space shuttle orbiter trajectory.‘4 By leadingedges for non-ablatinghypersonic operation. comparinga trajectory to an APC in this manner a One of the most useful approachesfor understanding vehicle designercan quickly determineif the vehicle the conditions wherethese leading edgesare capable will experienceleading edge ablation. In this caseno of operating without ablation is the aerothermal ablation occurs becausethe trajectory does not cross performanceconstraint (APC). APCs aredetermined the APC.

At high altitude the APC becomesnearly parallel to Mars wherea sharpleading edge will begin to ablate, ’ the ,velocity axis I (as shown by the, dashed andbeyond which the waveriderdesign should include extrapolation line) becausethe transition from the, a degreeof blowing, gas injected at the leading edge i continuum to the rarefied flow regime gradually and lifting the boundarylayer downstreamto simulate reducesthe aerothermalheating.3s Figure 9 shows ablation. For the ablating -leading edge case the typical AGA flight profiles for Venus and Mars, at upper’limit to the AGA speed,if a chine-typeleading equivalentEarth altitude, indicating that thereis some edgeablator is used,seems to be considerablyabove crossing of the APC. .Aero-gravity assist trajectories the value of V = 18 km/s computedhere. that cross the APC causeabiation of the sharp LJHTC leadingedge introducing significant uncertainty into It is clear’thatrapid small changesin vehicle angle of estimates of thermal protection system (TPS) attack’ are needed, responding to: measured lift performance, vehicle aerodynamics, and control accelerations,and analysesand tests are requiredto capability. Trajectories that do .not cross the APC provide the necessaryvehicle designparameters such ’ minimize theseuncertainties and simplify the vehicle as control surfacearea and location, centerof gravity ’ designprocess. By utilizing APCs in this mannera position, andpitch and yaw stability. Heating’is a vehicle designerwould be ableto quickly determineif major parameterbut being confined mainly to the the currentvehicle designwill experienceleading edge leadingedge it can be dealt with. ablation .without waiting for an extensive thermal analysis. Simplifying the design process is _’ important when attempting t-o rapidly move forward / from conceptsto actual flight hardware. Tests such RECOMMENDATIONS as in the SHARP program will help to, further understandand calibrate the APC anaIysesand the 1. System studies in more detail are requiredfor applicationto the leadingedges of the AGA waverider severalcandidate missions. Two main classesof such vehicles. missions are (1) atmospheric sample return from ., Mars, with AGA at Mars, fbr isotopic analysis, and Other technology developmentneeds include control the same for Venus and other planets, including surface design, control computational design and severalplanets on one flight: and (2) transportation software, and avionics hardwaredevelopment. A type flights to Mars and other planets, delivering an program will be necessaryto develop and.fly test orbiter or lander via a ballute, performing AGA to models of hypersonic waveriders subjected to return to Earth orbit for the next payload. environmentssimilar to the AGA maneuversat the planets. Such a program is in the planning stageby 2. Studies of rapid vehicle lift control in AGA, and the AGA teamrepresented in this paper. development of an adaptive exit algorithm from measured drag and the estimation of the computationalperformance,that will be necessaryon- CONCLUSIONS board.toexecute the algorithm.

In general,AGA enhancesthe gravity-assisttechnique 3. Waveriderdesigns including leadingedge blowing, tremendously by reducing launch energy and and appropriatewind tunnel simulations. decreasingflight time. This enhancementdepends on the availability of high lift-to-drag hypersonic 4. Earth entry waveridertests from secondarypayloads vehicles, as exemplified by current literature on the on Delta or Ariane vehicles, entering.at steep angles waverider. The results demonstratethat AGA is an to give strong heating and ablation, and verification enabling technology that can significantly reduce of aerocharacteristics. mission costs, increase science return, and allow - greateraccess to the Solar System. 5. Aidevelopmentprogram for ultra-high temperature non-ablating materials, charring ablator and low- From this approximateanalysis one can concludethat density.high temperatureinsulating materials should be actively pursued. a vehicle with a mass about lo-15 kg/m2 of plan area ‘% can be designedto fly AGA at Venus and Mars, beginning at Earth with a C3 of up to 50 km2/s2. There is a limiting, speedyV in AGA at Venus,and *’ _

10 ._ ., (c)l999 American Institute of Aeroriautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. l

Y I REFERENCES 11. Randolph,J. E., McRonald, A:D., Solar System Fast Mission Trajectories Using Aero- 1.. Randolph, J. E., Aero-Gravity-Assist: Gi%ity-Assist, AIAA PaperNo. 91-0531, 7 January TrajectoryAnal,y&s for Starprobe,JPL Internal IOM 1991,and J. Spacecraftand Rockets, Vol. 29, No. 2, 312182.5-981,5 August 1982 pp. 223-232,March-April,, 1992.

2. Longuski, J. M., Can Aero-Gravity-Assist 12. Lewis, M. J., McRonald, A. D., Design of through the Venusian Atmosphere Permit a Near HypersonicWaveriders for. AeroassistedInterplanetary Radial Trajectory into the Sun, JPL EM’ 312/82-133, Trajectories,AIAA Paper No. 91-0053, 7 January December1982 . 1991

3. Nonweiler, T. R. F., AerodynamicProblems 13. Gillum, M., Kammeyer,M.+ Burnett, D., of Manned Space Vehicles, Journal of the Royal Wind Tunnel Results for a Mach 14 Waverider, AeronauticsSociety, Vol. 63, September1959, pp. AIAA PaperNo. 94-0384,.lO January1994 521-528 14. Williams, S. N., Automated Design of 4. Bowcutt, K. G., Anderson, J. D., and Multiple Encounter Gravity-Assist Trajectories, Capriotti, D., Viscous Optimized Hypersonic Mastersthesis, Purdue University, August 1990 Waveriders,AIAA PaperNo. 87-0272, 12 January 1987 16. : Takashima,N., and Lewis, M.J., “‘A Cone- WedgeWaverider Configuration for Engine-Airframe 5. Anderson,J. D., Jr., A Survey of Modem Integration,”AIAA Journal of Aircraft, Vol. 32, No. Researchin Hypersonic Aerodynamics,AIAA Paper _ 5, pp. 1142-144,Sept-Oct. 1995. No. 84-1578,25 June 1984 17. Burnett, D., and Lewis, M.J., “A 6. Randolph, J. E., McRonald, A. D., Solar Reevaluation of the Waverider Design Process,” ProbeMission Status, AAS Paper89-212, 24 April AIAA 93-0404presented at AIAA AerospaceSciences 1989 Meeting, Reno,Nev., Jan. 1993.

7. Lewis, M. J., Kathari, A. P., The.Use of 18: Tarpley, C., and Lewis, M.J., “Stability and Hypersonic Waveridersfor Planetary-Exploration, , Control of Hypersonic WaveriderVehicles, ” AIAA presented at the AIAA/JPL 2nd International Journal of Aircraft, Vol. 32, No. 4, pp. 795-803, Conferenceon Solar System Exploration, 22 August JulyLAugust1995. 1989 19. Takashima, N., and Lewis, M. J. 8. Lewis, M. J., The Use of Hypersonic “Optimization of Waverider-BasedCruise Vehicles Waveridersfor Aero-AssistedOrbital Maneuvering, with Off-Design Considerations,“AIAA Journal of Proceedingof the 30th InternationalConference on Aircraft, Vol. 36, No. 1, Jan-Feb1999, pp. 235-245. Aviation and Space,February 1990 20. .J.M. Ldnguski and S.N. Williams, 9. McRonald, A. D., Randolph, J. E., “AutomatedDesign of Gravity-Assist Trajectoriesto Applications of Aero-Gravity-Assist‘to High Energy Mars andthe Outer Planets,”Celestial Mechanicsand Solar System Missions, AIAA PaperNo. 90-2891, Dynamical Astronomy, Vol. 52, No. 3 pp. 207-220, 20 August 1990 1991.

10. McRonald, A. D., Randolph, J. E., 21. E.P. Bonfiglio, J.M. Longuski, and N.X. Hypersonic .Maneuvering to Provide Planetary Vinh, “Automated Design of Aerogravity-Assist Gravity Assist, AIAA PaperNo. 90-0539, 8 January Trajectories,” AAS 99-361, Anchorage Alaska, 1990 August 1999, AASIAIAA AstrodynamicsSpecialist Conference.

11 (c)l999 American Institute of Aeronautics 81Astronautics or published with permission cif author(s) and/or author(s)’ sponsoring organization.

t. 22. E.P. Bonfiglio, “Automated Design of 33. I P.Kolodziej, J.D. Bull, F.S. Milos and T.H.. ’ Gravity-Assist and Aerogravity-Assist Trajectories,” Squire, Aerothermal Performance Constraints for Master’s Thesis, School of Aeronautics and: Small, Leading Edges Operating at Hypervelocity, Astronautics, PurdueUniversity, ’ West Lafayette IN, NASA CP-3359,91-98, Sept. 1,997. August 1999. 34. j P.Kolodziej,J.Bull,J,Salute, and D.L.Keese,’ 23 J.A. Sims, J:M. Longuski, andM.R. Pate& First Flight Demonstration,of a Sharp Ultra-High “Aerogravity-Assist Trajectoriesto the Outer Planets TemperatureCeramic, Nosetip, NASA TM: I 12215, and the Effect of Drag,” To Appear in Journal of Dec. 1997. Spacecraftand Rockets. 35. P. Kolodziej, Aerothermal Performance 24. J.M. Ludwinski, “Pluto-Kuiper Express Constraintsfor Hypervelocity Small Radius Unswept ‘Preliminary Mission Plan,” Jet Propulsion Leading Edges and Nosetips, NASA TM 112204, Laboratory, JPL Publication, Pasadena CA, July 1997. December1998.

25. M.R. Pate], “AutomatedDesign of Delta-V Gravity-Assist Trajectories for Solar System Exploration,!’ Master’s Thesis, SchooI of Aeronautics and Astronautics, PurdueUniversity, West Lafayette IN, August 1993.

26. M.R. Patel, J.M. Longuski, and J.A. Sims, “Mars FreeReturn Trajectories, ”Journal of Spacecraft and Rockets, Vol. 35, No. 3, pp. 350-354,May-June 1998.

27. (Ed.) A. Kliore, The. Venus International ReferenceAtmosphere,Advances in Space Research, Vol. 5, No. 11, PergamonPress, 1985.

28. (Ed. ) A. Kliore, The Mars Reference Atmosphere, Advancesin Space Research,Vol. 2, No. 2, PergamonPress, 1982. . I 29. t L. L. Perini, Compilation and Correlation of Stagnation Convective Heating Rates on Spherical Bodies, J. Spacecraft& Rockets, Vol. 12, March, 1975.

30. W.A. Page and H.T. Woodward,Radiative and Convective Heating During Venus Entry, AIAA J., Vol. 10, No. 10, 1379-81,October, 1972.

31. K. Sutton .and R. A. Graves, “A general Stagnation-Point Convective Heating Equation for Arbitrary Gas Mixtures” NASA TR R-376, November 1971.

32. J.N.Moss and E.C. Anderson;Aerothermal Environment for Jupiter Entry Probes,? 1976.

_’ 12 (c)l999 American Institute of Aeronautics & Astronautics or published with-permission of author(s) and/or author(s)’ sponsoring organization. 8

Angle of Attack Data For Mach 60 Waverider Optimized u Fqr Maximum L / D in Venus Atmosphere 6 M.-. /’ -.t_ -.-.-.- L,D /’ -.-. 0.15 ,6 b--l ----c i -, CL i 0.125 4b1- -. c, 1 !‘. ‘. /:

22; .:. .,, ._. i j 0.1 B _ i 2 - :' 2020: .i. _a- - 0.075 4 Q I i Oh g -27-2 i i / - 0.05 P I I' L? i / i .*.:’ i ,' *.’ i I’ i ,/ c , -,/ ’ -0.025 ”

-4 -2 0 2 4 a.Venture Star b.Deita II 7925H Angle of Attack (Degrees) _. Fig. 1. Illustration of waverider : (a) carried to 2(b) ~~erodynamic perfOmaUCer turbu1ant boundsry LEO on the Venture Star; (b) carried to LEO on leyar; the Delta II 7925H

Planform View Planform View

5 Side View

Base View

Isometric View Base View

rig. 2. (a) Waverider 8hapm for a turbultit 2(c) ebape for bmhar bound- lw=r; boundary layer

V 70.14

APPROACH DURING DEPARTURE .’ / c AGA 2I- .../. -... /.../. 4 m6,p i i / / / 0.040; / i / / -1n ,i* , J" n s-&n

a (Degrees) : APPROACH bM&tNG DEPARTURE Z(d) Aarac! pufo rpmcmfora laminarboupdary layw ,.

_- rig. 4. Illumtratioa of AQA: 'rotation and dacrbase of'the Vinf vector; and trajectory options

: - .rig. 3. Travel times for Pbto ogportuniths with Mu at Venus and Mars. (c)l999 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization.

1 . I I I I I 2 4 6. 6 10 12 14 16 EARlHUUNCHVELOt2WVhfE.blh I I I ---_ I 2 4 6 6 10 rig. 5. Approximate parzunektsrs for AGA at Vuru8 EARTH LAUNCH VELOCrrY V ~&nb ad uars ate a function of Vinf at Earth:

5(a) Planet velocities 5(c) Atmompheric flight time .

101 -a 8 . !!I

P P 6 2 *50 \I.

EARTH LAUNCH VELOCIW. V k(tkmfs

0 5(d) Interplanetary traxk time8 5 0 10 20 3 vl”#Dv lntv bn+s 5(b) turn angles (c)l999 American Institute of Aeronautics & Astronautics qr published with permission :of author(s) and/or author(s)’ sponsoring organization.

k 12 I Ial ’ ’ (b 180 190 200 VENUS EXIT ANGLE. dq

2 2 9 El

$ > 9 7 4’ !i 10 h I I 3 . I NREULEI uf 7

PI 6 -a . c9

5 02 0.3 0.4 0.5 0.6 loo 120 140 160 164 TIME FROM VENUS TO hiARS. YEAR VENUS-MARS TRAVEL RANGE ANGLE.dw I I I I 1I - 2 4 6 I. -2 0 I_-. - -- --. ANGLE OF ATTACK. dW Tie. 8. VO~~-H~~S trmol tin (a) d trwol ruwi mg10 (b) 0~ a fun&ion of vi&v for different dt u@*s at vmz4um. Fig. 6. (a) Bl= m/CLA and (b) L/D as a function of angle of attack (AOA).

450 -

I 4@4l , T;PICAL Ah lWdClORlES I I I I II II I

. I .

., -: rig. s., C~a&on*bet&&l a Shuttle trajectory' and ri&-Bl aerothermal perfornrsncs conPtr8int. rag. 5. l$Quilibrium ~~.&mospharic density and omv@ctive heating rats am a function of 81' = m/CLh ., for V - 12 to 19 k20m/8 in ma at Venus “1 Mars.