American Institute of Aeronautics and Astronautics1
Ballute Aerocapture Trajectories at Neptune
Daniel T. Lyons* and Wyatt R. Johnson.*
California Institute of Technology, Jet Propulsion Laboratory, Pasadena, CA, 91109
Using an inflatable ballute system for aerocapture at planets and moons withatmospheres has the potential to provide significant performance benefits compared not onlyto traditional all propulsive capture, but also to aeroshell based aerocapture technologies.This paper discusses the characteristics of entry trajectories for ballute aerocapture atNeptune. These trajectories are the first steps in a larger systems analysis effort that isunderway to characterize and optimize the performance of a ballute aerocapture system forfuture missions not only at Neptune, but also the other bodies with atmospheres.
I. IntroductionThe name "Ballute" is a cross between a "balloon" and a "parachute". The inflated components provide the
stiffness needed to deploy the structure in a vacuum and then maintain the proper shape of a very light weightstructure while generating sufficient atmospheric drag to capture the vehicle into orbit. The large drag area acts likea parachute to slow the spacecraft rapidly once it enters the upper atmosphere of the target body. Preliminary studiesof ballutes for aerocapture at several planetary bodies were pioneered by Angus McRonald.1,2,3 Jeff Hall has alsomade recent contributions to the advancement of ballute technology.4 Currently, the In Space Propulsion program isfunding an interdisciplinary team of engineers lead by Kevin Miller and Jim Masciarelli (Ball Aerospace). This teamis taking a closer look at characterizing and refining the use of ballutes for future aerocapture missions.5-9 The teamincludes experts from Ball Aerospace (system engineering, thermal, structures, control, and management), ILCDover (inflatable structures), NASA Langley Research Center (aerothermodynamics and hypersonic performanceverification and wind tunnel testing), and the Jet Propulsion Laboratory (trajectories, mission design, andinstrumentation). Preliminary calculations have shown that Titan aerocapture ballutes could be constructed usingexisting materials such as Kapton or Upilex. These large, lightweight inflatable structures would provide asignificant mass savings over traditional all-propulsive vehicles. A ballute even provides a significant mass savingwhen compared to aerocapture using an aeroshell. The mass savings are even larger if the aeroshell design includesa special transfer stage to provide power and attitude control during cruise. In addition to the low additional mass ofthe ballute for aerocapture, one of the fundamental benefits of carrying a ballute is that the primary spacecraft busdoes not have to remain tightly packed inside the aeroshell during cruise, but can be deployed and flown like anorbiter during the long cruise phase from Earth to Neptune.
All propulsive capture requires that the spacecraft must carry all of the propellant needed for the mission. Forlow altitude orbiters, the mass of the propellant for a traditional all propulsive spacecraft becomes so large that theuseful science payload becomes too small to be cost effective. In some cases, such as missions to Titan and Neptune,it may not be possible to conduct an orbital mission without aerocapture and/or other advanced propulsiontechnologies. One alternative for reducing the amount of propellant that must be carried is to use atmospheric dragto provide the velocity change required to capture into orbit.
A. Ballute Aerocapture Basics: High Drag, Low Heating
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Imagine the approach used for aerobraking, where the spacecraft is so high in the atmosphere that the heatingrate is tolerable even for an unprotected spacecraft. For a given initial entry orbit, if the area of such a high altitudespacecraft is increased, the amount of drag force on the spacecraft increases, but the heating per unit area remainsrelatively constant. The ballute concept takes this idea to the limit by dramatically increasing the area of thespacecraft. Enough drag is produced by the very large area of the ballute to remove the energy required to captureinto orbit in a single pass through the atmosphere, while the vehicle remains high in the atmosphere where theheating rates are relatively low. Assuming that the entry velocity is determined by the interplanetary trajectory, theheating rate is primarily a function of the atmospheric density, but the drag force is a function of both the densityand the projected frontal area. The ballute system can be designed so that an unprotected spacecraft could survivethe aerocapture heating rates if the drag-producing area is made large enough. If the thermal limits are set at about500°C by the thin film Ballute material, such as Kapton, then the spacecraft bus would require a thin thermal blanketto protect the exposed surface during the peak heating pulse. If the thermal limits are set low enough for anunblanketed spacecraft, then a much larger ballute is needed. The system design of the specific project would haveto balance the size and mass of the ballute against the mass of the thermal protection and mission objectives of theprimary spacecraft to optimize the design.
B. Ballute Team History
The work described in this paper is a very small part of a much larger on-going effort to increase the TechnicalReadiness Level of ballute technology. The team was formed in the last quarter of 2000 by Kevin Miller (BallAerospace) for a proposal to the Gossamer program. The team was a consortium of individual specialists fromIndustry, Academia, and NASA. Ball Aerospace has been studying the spacecraft system, including thermalanalysis, structures, control, and mass budgets. Dick Wilmoth at NASA Langley was the coinvestigator in charge ofcomputing the flow field around the ballute/spacecraft system. Peter Gnoffo took the lead on the computational flowanalysis when Dick retired. Professor Jim McDaniel was subcontracted through Langley to perform wind tunneltests in a special facility he runs at the University of Virginia. Ballute models are also being tested in the NASALangley 20 inch Mach 6 Air and CF4 wind tunnels. Jim Stein is leading a materials test and fabrication study at ILCDover to evaluate various candidate ballute materials and construction techniques. The focus of the study for theGossamer program was for aerocapture at Mars using a towed ballute. Although the Gossamer program was phasedout, the team was able to win an award from the In Space Propulsion program to study aerocapture at Titan andNeptune. Titan aerocapture using a trailing ballute concept was studied extensively in 20036-8. The In SpacePropulsion program funded a second study of an attached ballute concept in 2003, however, budget uncertaintiesdelayed the start of that study until this year. The co-principal Investigator for the attached ballute study is JimMasciarelli. The current emphasis is to bring the attached ballute concept9 to the same level as the trailing ballutestudy by the end of the 2004 fiscal year. Although ballute technology is in constant danger of “infant mortality”when competing for funding against technology that has a more immediate payoff, the impressive progress that hasbeen made by our team has been enough to keep ballute technology development alive.
II. Neptune ResultsThe trajectories described in this paper are a very preliminary analysis that will be used to size the ballute that
will be used for more detailed analyses. A simple ballistic transfer from Earth to Neptune was used to characterizethe arrival Vinfinity that might be reasonably expected at Neptune. Then a series of ballute aerocapture trajectorieswere generated for several different values of ballute area for a range of arrival Vinfinities (entry speeds) that spannedrange of probable values. An aeroshell aerocapture study for Neptune10 that was conducted by the In SpacePropulsion Program provided a starting mission scenario, including the final target orbit parameters.
C. Arrival Vinfinity Characterization
Figure 1 shows the arrival Vinfinity (the lower, red curve) and the launch Vinfinity (upper blue “stalactites”) versusthe flight time in years for an arbitrary fixed arrival date of Dec. 29, 2049. Although the aeroshell reference missionwas based on a low thrust trajectory with gravity assists from the inner planets, the arrival Vinfinity will be in the samerange as the ballistic trajectory – especially if the last flyby is at Earth, rather than Venus – but the flight time will beseveral years longer than for the purely ballistic trajectory shown here. The figure shows that really short flight times
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have very high arrival speeds. Similarly, extremely long flight times can reduce the arrival Vinfinity to about 4 km/sec.Unfortunately, extremely long flight times drive up the operations cost for the mission, while extremely short flighttimes result in extremely high heating, so the design goal is to find the appropriate design point that will minimizethe total cost of the mission.
Figure 1. Launch and Arrival Vinfinity versus Flight Time for a fixed Arrival Date.
Figure 2 shows the Entry Speed versus the Arrival Vinfinity. Two curves are shown for two different referencealtitudes. The 1000 km reference altitude is believed to represent the entry altitude used in the aeroshell analysis.The 4000 km altitude is the highest altitude for which the density is defined in the NeptuneGRAM atmosphericmodel that was developed by Jere Justus for the Neptune aeroshell study that was sponsored by the In SpacePropulsion Program. The ballute team will pick an altitude between these two values to use as the Entry Altitude forthe ballute mission. Later discussion will show that 4000 km is well above the altitude at which drag becomes afactor for ballute entry, but that 1000 km is a little below the point where noticeable drag begins. Some of the laterfigures reference entry to 4000 km because that was the top of the NeptuneGRAM atmosphere. A better value forentry altitude is discussed later in this paper. Using the arrival Vinfinity as the reference quantity eliminates theconfusion associated with picking a particular entry altitude.
Figure 2. Entry Speed versus Arrival Vinfinity for Two Candidate Entry Altitudes
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The aeroshell reference mission targeted a highly eccentric final orbit with an apoapsis altitude of 430,000 km.The eccentric orbit was selected to provide close, periodic flybys of the moon Triton. Since the energy of the targetorbit at atmospheric exit is nearly constant, the spacecraft must leave the atmosphere with a particular speed for allcases. The entry speed at that reference altitude is determined by the arrival Vinfinity. The difference between the entryand exit speeds at a given reference altitude represents the change in speed that was caused by atmospheric drag.Figure 3 shows that for the lowest arrival Vinfinity under consideration (4 km/s), a change of only 1 km/sec is neededto capture into the desired orbit, while a 6 km/s change is required for an arrival Vinfinity of 16 km/s.
Figure 3. Change in Velocity to Achieve 430,000 km Apoapsis vs Arrival Vinfinity
Based on the ballistic transfer from Earth to Neptune, arrival Vinfinities between 4 and 16 km/s were selected forevaluation as part of the ballute study. Vinfinity = 4 km/sec requires an extremely long transfer time from Earth toNeptune! Vinfinity = 16 km/sec represents such a short transfer time, that the departure Vinfinity becomes extremelyhigh for the ballistic transfer. The low-thrust baseline developed for the aeroshell study has an arrival Vinfinity similarto that at the high end of the range studied, although the aeroshell used a low-thrust vehicle and several gravityassists from the inner planets to minimize the launch costs. A ballistic transfer from Earth to Neptune with the same10 year cruise duration as the low-thrust aeroshell trajectory would have a lower arrival speed (Vinfinity ≈ 12 km/secrather than 16 km/sec), but would require a larger launch vehicle for the same payload. A 17 year cruise would beneeded to lower the arrival Vinfinity to ≈ 6 km/s.
D. Ballute Entry Trajectory Simulations at Neptune.
The preliminary ballute entry trajectory simulations used the following parameters. The entry mass was 500 kgfor all cases to make it easier to compare to the preliminary analysis that was run at Titan. Most cases use an area of1477 m2, because the first guess at a possible entry speed needed this area to achieve a reasonable heating rate. Twoother areas (750 m2 & 3000 m2 ) are evaluated to show the sensitivity of the trajectory to other areas for the sameentry mass. A constant drag coefficient = 1.37 was used for the ballute, because this was a reasonable value to use atTitan. Note that the detailed flowfield analysis by the NASA Langley flow computations shows that both the sizeand configuration of the ballute system, as well as the instantaneous Reynolds and Knuden numbers play animportant role in the actual drag coefficient. The ballute never separates in these preliminary trajectories, so the S/CCD is not an issue, although it will be for more detailed analyses that are planned using increasingly sophisticatedmodels. The Target Apoapsis (430,000 km from the aeroshell study) is achieved by searching for the periapsisaltitude of approach hyperbola that results in a trajectory with an osculating apoapsis altitude at atmospheric exitequal to the target value.
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Figure 4 shows the maximum free stream heat flux, Qdot (0.5 Rho•V3, W/cm2) versus arrival Vinfinity (lowercurve). Qdot increases rapidly as the arrival Vinfinity is increased. Plotted on the same graph is the Entry speed in(km/sec – using the numerical Y-axis) at a reference altitude of 4000 km (upper curve). Later plots will show thatthe reference altitude could be lowered.
At least 3 different heating rates are needed to characterize ballute aerocapture with a towed balluteconfiguration because there are three different characteristic sizes. The free stream heat flux, Qdot, shown inFigure 4 is appropriate for nearly free molecular flow, which is the case for the thin tethers that would be used toconnect the large inflated towed ballute with the main spacecraft. Qdot is representative when the most of theatmospheric molecules are likely to hit the spacecraft without interacting with molecules that have alreadytransferred their energy and momentum to the spacecraft. The fraction of Qdot that is transferred to heat thespacecraft is higher for free molecular flow, because individual molecules can hit the spacecraft at full orbital speed.Although the inflated ballute is flying through the same atmosphere as the rest of the spacecraft, the ballute is ordersof magnitude larger than the tether, and can reach conditions that are best characterized by continuum flowapproximations where the heating is proportional to the square root of the density rather than the density itself. Thecharacteristic size of the spacecraft is in between these extremes, and so is the heating rate. Aeroshell designershave to be concerned not only with the instantaneous heating rate, but also with the total amount of heat that isabsorbed and then stored in the vehicle structure. Total heat load is not an issue for the ballute, because it is so lowin mass and large in area that it reaches the equilibrium temperature very quickly. The equilibrium temperaturebalances the incoming heating rate with radiative cooling at the same rate.
Figure 5 shows the Maximum Qdot (W/cm2) versus arrival Vinfinity for several ballute sizes. Figure 6 shows thesame data plotted versus Entry speed for a 4000 km reference altitude. Doubling the area essentially cuts Qdot inhalf. Halving the area doubles Qdot.
Figure 7 shows Qdot as a function of time since the start of the simulation. Since the initial state is outside of theatmosphere, nothing “interesting” happens before about 700 sec, so the x-axis has been scaled to show the region ofinterest. All of these trajectories are for initial conditions which have been independently targeted for each case suchthat each case exits the atmosphere with a 430,000 km apoapsis altitude without releasing the ballute. Thus themaximum Qdot occurs near periapsis for each case. Once a ballute size is selected, the nominal entry trajectory willbe targeted lower in the atmosphere to accommodate Navigation and atmospheric uncertainties, and the peak Qdotwill increase and move earlier.7 The maximum values from Figure 7 are what was plotted in Figures 4-6.
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Figure 5. Maximum Qdot versus Vinfinity for Three Ballute Sizes.
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Figure 6. Maximum Qdot versus “Entry” Speed at 4000 km
In the legend,Smaller speed is VinfinityLarger speed is at 4000 km.
Figure 8. Altitude versus Time for a Range of Arrival Vinfinities
Figure 8 shows the altitude versus time since the start of the simulation. Plotting the Dynamic Pressure versusaltitude shows the relationship between the Entry Altitude definition and the force acting on the system. Using 1%of the Maximum for the 6 km/s (Vinfinity) case as the threshold shows that an altitude of about 1200 km might be acandidate for the definition of the Entry altitude. The forces acting on the vehicle are only 0.01% of the“Maximum” at an altitude of 1500 km. Thus a good definition of “Entry” for the ballute cases would be about 1500km. A reasonable person might even pick 2000 km as the entry altitude, to essentially eliminate drag forceuncertainty on the trajectory. ( I prefer to use the osculating periapsis radius and Vinfinity, rather than entry speed andentry flight path angle in part because there is no clear cut way to define the entry altitude.)
Figure 9. Dynamic Pressure versus Altitude near 1% of Max. for a Range of Arrival Vinfinities
Figure 10 shows the periapsis altitude versus Arrival Vinfinity. Faster arrival speeds and/or smaller ballute areasrequire the trajectory to be targeted lower in the atmosphere (equivalent to a steeper entry angle of attack). Each ofthe trajectories described in this section required an iterative search to find the periapsis altitude that resulted inachieving the target apoapsis altitude at atmospheric exit (without releasing the ballute) for the specified approachVinfinity.
American Institute of Aeronautics and Astronautics8
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Smaller Areameans targetdeeper in theatmosphere toget enoughdrag.
Figure 10. Periapsis Altitude versus Arrival Vinfinity for several Ballute Areas
Figures 11 and 12 show the time history of the velocity and deceleration of the cases for the 1477 m2 ballutearea. Higher arrival Vinfinity means higher speed at entry. Achieving a specific target apoapsis altitude at exit meansthat all trajectories exit the atmosphere at the same speed. Larger entry speeds require a larger change in velocitywhile in the atmosphere which results in higher heating for a given ballute size. Although the deceleration levels arehigher for the faster entry speeds, the highest entry speed under consideration has a g-load of only 3.5 g’s which isactually less than the 4.5 g’s that were considered acceptable as part of the Titan aerocapture study7. Since thesepreliminary Neptune cases do not release the ballute, the equivalent Titan case, where the spacecraft achieves thetarget orbit without releasing the ballute, only has a g-load of about 3.0 g’s. The g-load for the arrival targeting thataccommodates uncertainties at Neptune will be higher than for these preliminary cases – once the arrival is targetedlower in the atmosphere to accommodate Navigation and atmospheric uncertainties. Assuming that the Titan balluteconclusions that a 5 g deceleration could be accommodated by the ballute, then deceleration g-loads are not the issueat Neptune.
In the legend,Smaller speed is VinfinityLarger speed is at 4000 km.
Figure 13. Atmospheric Density versus Time for a Range of Arrival Vinfinities
Figure 13 shows the atmospheric density as a function of time for the 1477 m2 cases. Higher entry speeds requiretrajectories that dive deeper into the atmosphere, so the maximum atmospheric densities are higher. Note that thedensity is a smooth function of time. Future studies will introduce “noisy” atmosphere models to evaluateperformance in Monte Carlo studies of the separation algorithm.
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In the legend,Smaller speed is VinfinityLarger speed is at 4000 km.
Dynamic PressuresAre less than forTitan (at the same Qdot), so DynamicPressure is not aDriver for Neptune.
Figure 14. Dynamic Pressure versus Time for a Range of Arrival Vinfinities
Figure 14 shows the time history of the dynamic pressure. The dynamic pressure is typically less than for Titanentry, so a ballute system that was designed to withstand the dynamic pressures at Titan would also survive thepressure loads at Neptune.
E. Comparison of a Titan and a Neptune trajectory.
The following compares a single Titan ballute aerocapture trajectory with a single Neptune trajectory. Sincenone of the Neptune trajectories release the ballute, the Titan trajectory that achieved its target apoapsis altitude of1700 km without ballute release was selected. The ballute area of 750 m2 was a typical value used in the Titanstudy. Since the Titan study used a Vinfinity of about 6.0 km/s, and the Neptune study looked at a variety of values,the Neptune trajectory with a maximum value of Qdot closest to the Titan study was selected for comparison,because Qdot is a key design driver. Figure 15 shows the free stream heat flux, Qdot, versus time from theMaximum Deceleration. (Since the trajectories started at arbitrary times, the point of maximum deceleration seemedto be a good reference value for plotting results.) By a surprising coincidence, the Vinfinity for the Neptune trajectorywas also 6 km/s, although the entry speed ( at 1500 km ) was 23.6 km/s, almost 4 times larger than the entry speedof 6.4 km/s ( at 1025 km) for Titan. Even more surprising is the fact that the duration of the drag pass is nearlyidentical for both bodies, even though the diameter of Neptune is an order of magnitude larger than Titan ! Theshape of the Titan Qdot profile is more asymmetric than that for Neptune, because the Titan trajectory is targeted toa low altitude, circular orbit, while the Neptune trajectory is targeted to a highly elliptical (430,000 km apoapsisaltitude) orbit. The entry mass and drag coefficients are the same for both trajectories, but the surface area of theNeptune ballute is twice that of the Titan ballute.
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, 1477 m2, 430,000 km Apo.Titan: 6.4 km/s V
ENTRY, 750 m2, 1,700 km Apo.
Vinfinity
= 6.0 km/sConstant CD = 1.37 m2Entry Mass = 500 kg
Figure 15. Qdot versus Time for a Neptune and a Titan trajectory with Similar Maxima.
Since the entry speed of the Neptune vehicle is nearly 4 times larger than that of the Titan vehicle, and since thefree stream heat flux, Qdot ( which is 1/2 Density • V3 ) is nearly equal, then the density of the atmosphere for theNeptune case must be significantly less than that for Titan. Figure 16 shows the atmospheric densities plotted versusthe time. Note that an exponential scale had to be used, because the maximum densities are different by two ordersof magnitude. The entry speeds, shown at the left in Figure 17, are a factor of 4 different, which cubed gives anorder of magnitude difference at entry. The Titan vehicle slows significantly, so that by the time it reachesperiapsis, where the density is highest, the velocity has been cut in half, which when cubed leads to another factor of10. The velocity of the Neptune vehicle is not changed nearly as much as a percentage of the entry speed.Combining these effects leads to two orders of magnitude difference in the maximum densities. The maximumdeceleration (used as the reference time) is close to periapsis for the Neptune trajectory, but occurs nearly 100 secbefore periapsis for the Titan trajectory. Another interesting coincidence is that the Titan trajectory leaves theatmosphere at 725 m/s less than escape velocity, with the Neptune trajectory leaves the atmosphere about 627 m/sless than escape. Since the Neptune value is nearly the same as the Titan value, but is a much larger percentage ofthe ∆V from the atmosphere, I expect acceptable performance from the separation algorithm7,10 once we have thechance to evaluate it at Neptune.
0.0010.010.1
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, 1477 m2, 430,000 km Apo.Titan: 6.4 km/s V
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Vinfinity
= 6.0 km/sConstant CD = 1.37 m2Entry Mass = 500 kg
Figure 16. Atmospheric Density for a Neptune and a Titan Trajectory with Similar Qdot.
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ENTRY, 750 m2, 1,700 km Apo.
Vinfinity
= 6.0 km/sConstant CD = 1.37 m2Entry Mass = 500 kg
Neptune ∆V = 1.325 km/s @ 1500 km Entry/Exit
Titan ∆V = 4.897 km/s @ 1025 km Entry/Exit
Figure 17. Speed Relative to the Atmosphere for a Neptune and a Titan Trajectory with Similar Qdot.
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ENTRY, 750 m2, 1,700 km Apo.
Vinfinity
= 6.0 km/sConstant CD = 1.37 m2Entry Mass = 500 kg
Figure 18. Deceleration Profiles for a Neptune and a Titan Trajectory with Similar Qdot.
Figure 18 shows the deceleration versus time for the Titan and Neptune example trajectories. As pointed outearlier, the deceleration acting on a Neptune vehicle tends to be less than that for a Titan vehicle. The difference islargely due to the differences in the target orbits. Nearly 4,900 m/s are removed by the Titan atmosphere, while only1,325 m/s are removed at Neptune. The required ∆V is detemined by the difference between the velocities at entryand exit. The entry velocity is determined by the interplanetary trajectory for a ballistic transfer, while the exitvelocity is determined by the target orbit. The Titan mission scenario called for a low circular science orbit thatrequired a large change in ∆V, while the Neptune mission scenario called for a highly elliptical science orbit.(Another 6 km/s would be required for a Neptune aerocapture into a low circular orbit … a scenario that mightrequire multiple passes through the atmosphere after capture.) Since the required ∆V is the integral under thedeceleration curve, and since the duration of the deceleration is similar, it is not surprising that the Titan decelerationis larger, at least for these two mission scenarios.
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III. Conclusions
For ballute aerocapture at Neptune, heating is to be the limiting factor. Reasonable heating is possible forballistic trajectories with long interplanetary flight times. The aeroshell reference study used low thrustinterplanetary cruise with multiple planetary flybys to try to minimize flight time without regard to the entry speed.Since ballutes benefit from low entry speeds, searching for interplanetary trajectories that try to minimize both theflight time and the arrival speed will be required.
This paper has illustrated the very preliminary ballute aerocapture trajectories needed to size the ballute forNeptune. The final ballute size and configuration is a tradeoff between mass and size of the ballute, the mass ofextra thermal protection, if any, for the spacecraft, and arrival speeds, which require additional operational costs.Previously reported Titan results7 are more detailed because more resources were available. Similar results forNeptune will be available once the funding becomes available.
These preliminary trajectory results are part of a system wide trade study which indicate that ballutes appearto be a feasible option for aerocapture at Titan and Neptune. The mass of the ballute required to achievetemperatures that are survivable with only minimal thermal shielding of the spacecraft is significantly less than themass of an aeroshell for the same system mass at atmospheric entry. A detailed mass comparison by the Ball teammembers showed that a ballute system mass is about 10% of the entry mass, while the equivalent aeroshell systemmass is more than 40% of the entry mass for aerocapture at Titan. Ballute aerocapture systems have the potential tosignificantly increase the mass available for a scientific payload for a given mission to Titan or Neptune. Thetremendous mass saving, coupled with very positive developments from the current ballute studies, make furtherinvestment in ballute technology highly desirable.
Acknowledgements
The work described was performed at the Jet Propulsion Laboratory (JPL), California Institute of Technologyunder contract with National Aeronautics and Space Administration (NASA). Funding for this study was providedby the NASA Office of Space Science through the In Space Propulsion Program. The authors would particularly liketo thank the trailing ballute Principal Investigator, Kevin Miller (Ball Aerospace) as well as Bonnie James, MichelleMunk, and Erin Richardson at the NASA Marshall Space Flight Center, In Space Propulsion Program office fortheir sponsorship and excellent guidance in the management of these efforts.
References
1. Angus McRonald, "A Light-Weight Inflatable Hypersonic Drag Device for Planetary Entry", AssociationAeronautique de France Conf. at Arcachon, France, March 16-18, 1999.
2. Angus McRonald, "A Light-Weight Inflatable Hypersonic Drag Device for Venus Entry" AAS/AIAAAstrodynamics Specialist Conf., Girdwood, Alaska, Aug 16-19, 1999. AAS 99-355
3. Angus Mc Ronald, “A Light-Weight Hypersonic Inflatable Drag Device For a Neptune Orbiter”, AAS/AIAASpace Flight Mechanics Meeting, Clearwater, FL, Jan. 23-26, 2000. AAS 00-170
4. Jeffery L. Hall , “A Review of Ballute Technology for Planetary Aerocapture”, 4th IAA Conference on LowCost Planetary Missions, Laurel, MD, May 2-5, 2000.
5. Kevin Miller, “Gossamer Ballute Aerocapture Final Report”. Sept. 25, 2002. Contract #1205966.6. Douglas Gulick, Kevin Miller, Jake Lewis, Bill Trochman, George Sapna, Jim Stein, Daniel T. Lyons, Richard
Wilmoth “Trailing Ballute Aerocapture: Concept and Feasibility Assessment“, 39th AIAA/ASME/SAE/ ASEE JointPropulsion Conference and Exhibit, Von Braun Center, Huntsville, Alabama. July 20-23, 2003.
7. Daniel Lyons, Wyatt Johnson, “Ballute Aerocapture Trajectories at Titan”, AIAA/AAS AstrodynamicsConference, Big Sky, Montana, August 2003. AAS 03-646.
8. Wyatt Johnson, Daniel Lyons, “Titan Ballute Aerocapture Using a Perturbed TitanGRAM Model”, AIAAAtmospheric Flight Mechanics Conference, Providence, Rhode Island, August 16-19, 2004. AIAA 2004-5280.
9. Peter A. Gnoffo, Brian P. Anderson,: "Computational Analysis of Towed Ballute Interactions," 8thAIAA/ASME Joint Thermophysics and Heat Transfer Conference, AIAA Paper 2002-2997, June 2002.
10. Rob Haw. “Aerocapture Navigation at Neptune”, AIAA/AAS Astrodynamics Conference, Big Sky, Montana,Aug. 3-7, 2003. AAS-03-643.
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