IPPW-13 Program Abstracts - Drag, Aerobraking and Aerocapture Session

Transcription

1 Future HIAD Advancements and Extension of Mission Applications HIAD on ULA (HULA) Orbital Reentry Flight Experiment Concept Drag, Aerobraking and Aerocapture Session Tuesday, June 14, 2016 from 9:15 AM to 1:30 PM Conveners: Al Witkowski Hannes Griebel Devin Kipp R. Keith Johnson 1 John DiNonno 2 SINGLE BODY IAD FOR MARS HIGH PAYLOAD MASS DELIVERY Maxim de Jong 3 RESULTS SUMMARY OF THE LOW-DENSITY SUPERSONIC DECELERATOR TECHNOLOGY DEVELOPMENT PROJECT Ian Clark and Mark Adler 4 ExoMars Parachute Systems: Design and Verification Steve Lingard Overview of the Parachute Decelerator System for the InSIGHT Mission Decelerator Pack Rotation Discovery During Qualification / Acceptance Testing for the InSight Mission PRELIMINARY STUDY ON MARS AEROCAPTURE AND LANDER MISSION USING DEEP SPACE MICRO SPACECRAFT AND MEMBRANE AEROSHELL Development of an Earth SmallSat Flight Test to Demonstrate Viability of Mars Aerocapture Al Witkowski Al Witkowski al.witkowski@zodiacaerospa ce.com al.witkowski@zodiacaerospa ce.com Kojiro Suzuki kjsuzuki@k.u-tokyo.ac.jp 8 Michael Werner mwerner9@gatech.edu 9 Results of the Venus Express Aerobraking Campaign Håkan Svedhem H.Svedhem@esa.int Improving Blunt Body Aerodynamic Performance Through Deployable Tabs Ashley M. Korzun ashley.m.korzun@nasa.gov 11 Use of Plasma Actuator to Improve Reentry Missions Sandra Coumar sandra.coumar@cnrsorleans.fr 12 July 13-17, 2016 Laurel MD USA 1

2 Future HIAD Advancements and Extension of Mission Applications R. K. Johnson 1, F.M. Cheatwood 1, A.C. Calomino 1, S.J. Hughes, G.T. Swanson 2, J.M. DiNonno 1, M.C. Lindell 1 1 NASA Langley Research Center, Hampton, VA, 23681, 2 AMA Incorporated, NASA Ames Research Center, Moffett Field, CA USA,, USA, 4NASA Ames Research Center, Moffett Field, CA USA Abstract: The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology has made significant advancements over the last decade with flight test demonstrations and ground development campaigns. The first generation (Gen 1) design and materials were flight tested with the successful third Inflatable Reentry Vehicle Experiment flight test of a 3 m HIAD (IRVE 3). Ground development efforts incorporated materials with higher thermal capabilities for the inflatable structure (IS) and flexible thermal protection system (F TPS) as a second generation (Gen 2) system. Current efforts and plans are focused on extending capabilities to improve overall system performance and reduce areal weight, as well as expand mission applicability. F TPS materials that offer greater thermal resistance, and packability to greater density, for a given thickness are being tested to demonstrated thermal performance benefits and manufacturability at flight relevant scale. IS materials and construction methods are being investigated to reduce mass, increase load capacities, and improve durability for packing. Previous HIAD systems focused on symmetric geometries using stacked torus construction. Flight simulations and trajectory analysis shows that symmetrical HIADs may provide L/D up to 0.25 via movable center of gravity (CG) offsets. HIAD capabilities can be greatly expanded to suit a broader range of mission applications with asymmetric shapes and/or modulating L/D. Various HIAD concepts are being developed to provide greater control to improve landing accuracy and reduce dependency upon propulsion systems during descent and landing. Concepts being studied include a canted stack torus design, control surfaces, and morphing configurations that allow the shape to be actively manipulated for flight control. This paper provides a summary of current HIAD development activities, and plans for future HIAD developments including advanced materials, improved construction techniques, and alternate geometry concepts that will greatly expand HIAD mission applications. July 13-17, 2016 Laurel MD USA 2

3 HIAD on ULA (HULA) Orbital Reentry Flight Experiment Concept J. M. DiNonno 1, F. M. Cheatwood 1, S. J. Hughes 1, M. M. Ragab 2, R. A. Dillman 1, R. J. Bodkin 1, C. H. Zumwalt 1, R. K. Johnson 1 1 NASA Langley Research Center, Hampton, VA, 23681, John.M.DiNonno@nasa.gov, F.M.Cheatwood@nasa.gov, Stephen.J.Hughes@nasa.gov, Robert.A.Dillman@nasa.gov, Richard.J.Bodkin@nasa.gov, Carlie.H.Zumwalt@nasa.gov, R.K.Johnson@nasa.gov, 2 United Launch Alliance, Centennial, CO 8011, Mohamed.M.Ragab@ulalaunch.com Abstract: This paper describes a proposed orbital velocity reentry flight test of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). The flight test builds upon ground development activities that continue to advance the materials, design, and manufacturing techniques for the inflatable structure and flexible thermal protection system (F-TPS) that comprise the inflatable heat shield. While certain aspects of material and system performance can be assessed using a variety of ground testing capabilities, only orbital velocity energy on a trajectory through the gradient density of the atmosphere can impart the combined aerodynamic and aeroheating design environments in real time. To achieve this at limited cost, the HIAD would be delivered to a spin-stablized entry trajectory as a secondary payload on the Centaur stage of a United Launch Alliance (ULA) Atlas V launch vehicle. Initial trajectory studies indicate that the combination of launch vehicle capability and achievable reentry vehicle ballistic numbers make this a strategic opportunity for technology development. This 4 to 6 meter diameter scale aeroshell flight, referred to as HIAD on ULA (HULA), would also contribute to ULA asset recovery development. ULA has proposed that a HIAD be utilized as part of the Sensible, Modular, Autonomous Return Technology (SMART) initiative to enable recovery of the Vulcan launch vehicle booster main engines [1], including a Mid-Air Recovery (MAR) to gently return these assets for reuse. Whereas HULA will attain valuable aerothermal and structural response data toward advancing HIAD technology, it may also provide a largest-to-date scaled flight test of the MAR operation, which in turn would allow the examination of a nearly pristine post-entry aeroshell. By utilizing infra-red camera imaging, HULA will also attain aft-side thermal response data, enhancing understanding of the aft side aerothermal environment, an area of high uncertainty. The aeroshell inflation will utilize a heritage design compressed gas system to minimize development costs. The data will be captured to both an onboard recorder and a recorder that is jettisoned and recovered separately from the reentry vehicle to mitigate risk. This paper provides an overview, including the architecture and flight concept of operations, for the proposed HULA flight experiment. References: [1] Ragab M. M., Cheatwood F. M., et. al. (2015) Launch Vehicle Recovery and Reuse, AIAA July 13-17, 2016 Laurel MD USA 3

4 SINGLE BODY IAD FOR MARS HIGH PAYLOAD MASS DELIVERY. M. de Jong 1, S. J. Saikia 2, E. Sheta 3, and V. Venugopalan 4, 1 Thin Red Line Aerospace Ltd., Chilliwack, Canada, maxim@thin-red-line.com, 2 Visiting Assistant Professor, ssaikia@purdue.edu, School of Aeronautics and Astronautics, Purdue University, West Lafayette, IN, 3 Senior Principal Engineer and Manager, efs@cfdrc.com, 4 Principal Engineer, vv@cfdrc.com, Aerospace & Defense Division, CFD Research Corporation, Huntsville, AL. Introduction and Challenge: NASA s 2014 Strategic Plan mandates a long-term goal to send humans to Mars, whereby the requisite associated technologies and capabilities are to be developed and demonstrated. Inflatable Aerodynamic Decelerator (IAD) technology is acknowledged to provide game changing capability in delivery of payload masses of tens of metric tons to Mars as well as to other celestial atmospheric bodies. However, in order to achieve the ballistic coefficient required to deliver large mass payloads of future exploration missions, deployable hypersonic decelerators will require deployed diameters in excess of 20 meters. This represents a significant scalability challenge, as current flight demonstrations have been limited to 3 m for an inflatable decelerator [1]. Many inflatable structural designs have produced ambiguous results due to unavoidable nonlinearities that manifest in overall performance predictability, scalability and flexural attributes. These problems are further exacerbated by complex multi-layered fabrication processes; nonlinear aeroelastic and aerothermal effects encountered during deceleration; and the need for multiple deceleration profiles to accommodate hypersonic, supersonic and subsonic flight regimes. Therefore, NASA is further investigating innovative IAD architectures and seeking fundamental studies of alternate guidance and control concepts that may be more amendable to implementation on large entry vehicles. Thin Red Line Single Body Spherecone: Thin Red Line Aerospace has advanced development and validation of a highly simplistic, scalable, mass-optimized, inflatable-deployable architecture for NASA that is capable of providing tailored performance for planetary entry vehicles. This Ultra High Performance Vessel (UHPV) technology (figure 1, left) was originally developed to help solve NASA s challenges with performance variation and excessive safety factor requirements for inflatable structures. Its fully determinate load paths allow unprecedented performance predictability regardless of scaling; as well as the lowest mass configuration of any competing design. UHPV can be fully characterized and validated analytically without the need for complex or extensive testing. Its proprietary planar fabric component design, enhances packaging efficiency and vastly lowers manufacturing costs. The UHPV Spherecone architecture (figure 1, center and right) comprises a single, primary inflatable volume of very low pressure yet exceptionally rigid and resistant to aeroload deflection. Subscale UHPV test articles can extensively validate 20 to 40 meter IAD behavioral attributes and performance. Transformable or Morphable Entry Systems: NASA seeks to Enable a vehicle to change shape or configuration to achieve additional functions during entry, descent, and landing, such as providing direct alpha and beta control or direct ballistic number control [1]. Current investigations show promise in UHPV shape morphing capability through manipulation of the meridional tendon array. Acknowledgement: This work was supported in part by NASA contracts NNX11CA28C Integrated Inflatable Ballute for Planetary Entry, NNL12AA04B Concept Development for a 10m Diameter Hypersonic Inflatable Aerodynamic Decelerator (HIAD), and NNX14CL24P Development of Hot Structures Materials for Inflatable Heat Shield. [1] NASA technology Roadmaps, TA 9: Entry, Descent, and Landing Systems. July, Figure 1. Left: Thin Red Line Ultra High Performance Vessel (UHPV). Center: UHPV Spherecone test article. Right: UHPV spherecone aero-mechanical analytical model July 13-17, 2016 Laurel MD USA 4

5 RESULTS SUMMARY OF THE LOW-DENSITY SUPERSONIC DECELERATOR TECHNOLOGY DEVELOPMENT PROJECT. Dr. Ian Clark 1 and Dr. Mark Adler 1, 1 NASA Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Dr., Pasadena, CA 91109, ian.g.clark@jpl.nasa.gov Abstract: In 2010 NASA initiated the Low-Density Supersonic Decelerator Project (LDSD) as one of the initial Technology Demonstration Missions in the Space Technology Mission Directorate. The purpose of LDSD was to develop the next generation of supersonic aerodynamic decelerators to significantly improve the landed mass, landed altitude, and landed accuracy capabilities of future Mars missions. Towards that end, LDSD began developing three new decelerator technologies, two Supersonic Inflatable Aerodynamic Decelerators (SIADs) and a new large-diameter supersonic parachute. A fourth decelerator technology, a large-diameter trailing ballute was also developed, originally as a means of parachute extraction. The first of the SIADs developed by LDSD, SIAD- R, was targeted at near term improvements in robotic mission capabilities and maximizing the landed mass capability of an Atlas V class launch vehicle. The 6 m diameter SIAD-R consists of a large, inflatable torus that is attached at the periphery of an entry vehicle and designed for deployment at Mach numbers of up to four. It was developed with the intention of acting as a rigid structure that would resist appreciable aeroelastic deformation during flight. The second SIAD developed, SIAD-E, is a larger, 8 m diameter decelerator that leveraged work performed in the 1960 s and 1970 s on an attached isotensoid concept. The size of the device was based on the desire to enable entry masses associated with the capability of a Delta IV-H class launch vehicle. SIAD-E is a predominantly ram-air inflated structure whose geometry is much more closely coupled with the oncoming flowfield than the rigid SIAD-R decelerator. As such, it was intended to represent the first forays into the development of a SIAD with aeroelastic coupling. In order to maximize the benefits associated with the higher Mach SIADs, a large supersonic parachute was also developed under LDSD. The 30.5 m diameter parachute was intended to improve on the performance of the heritage Disk-Gap-Band parachute not just in size but also in drag coefficient performance and extensibility to future missions. Due to the size of the devices and the energies necessary to test at relevant conditions, existing test infrastructure proved to be insufficient for the LDSD technologies. Thus, LDSD also developed a number of innovative approaches, including ground based rocketsled tests and high-altitude supersonic tests. This paper summarizes the techonlogies, test architectures, and test results from the six year LDSD effort. It will also describe the maturity level achieved for each decelerator technology and future directions for continuation of development activities. At the conclusion of 2015, LDSD had successfully matured two decelerator technologies, SIAD-R and a large supersonic ballute, to a maturity level appropriate for adoption by a Mars flight project. Two supersonic tests of a large 30.5 m parachute provided a significant amount of data on the inflation and operation of supersonic parachutes, but the failure of both parachutes during those tests exposed a deficiency in the understanding of the stresses and loads internal to the parachute canopy during supersonic inflatin. Lastly, ground testing of the 8 m SIAD-E article yielded mixed results and highlighted some of the difficulties associated with the testing of large, flexible aerodynamic decelerators. July 13-17, 2016 Laurel MD USA 5

6 . ExoMars Parachute Systems: Design and Verification J.S. Lingard 1, J.C. Underwood 1, S. Langlois 2 1 Vorticity Ltd, Chalgrove, Oxfordshire, OX44 7RW, United Kingdom, Steve.Lingard@vorticity-systems.com 2European Space Agency, Keplerlaan 1, 2201 AZ Noordwijk, Netherlands, Stephane.Langlois@esa.int Abstract: The key element for accessing the surface of Mars and one of the greatest challenges in space exploration is the successful execution of the entry, descent and landing sequence on the surface of the planet. Two missions are foreseen within the ExoMars programme: one consisting of the Trace Gas Orbiter plus an Entry, Descent and landing demonstrator Module (EDM), known as Schiaparelli, launched on 14 March 2016, and the other, featuring a rover, with a launch date of Both missions will be carried out in cooperation with Roscosmos. Although designed to demonstrate entry, descent and landing technologies, the ExoMars 2016 EDM also offers limited, but useful, science capabilities. The EDM will deliver a science package that will operate on the surface of Mars for a short duration after landing, planned to last approximately 9 days. As soon as the 600 kg EDM has slowed down to Mach 2, the parachute will be deployed to further decelerate to subsonic speed. A single stage parachute system has been selected. A mortar will deploy the 12.0 m diameter disk-gap-band scaled from the successful Huygens mission main parachute. The parachute must decelerate and stabilize the probe through the transonic regime and provide a stable, vertical trajectory above the landing site and a descent rate of circa 70 m/s. At Mach 0.4, the module will release the front heat shield. The EDM will activate its Doppler radar altimeter and velocimeter to locate its position with respect to the Martian surface. At the optimum distance from the ground, the lander will separate from the back shield and the liquid propulsion system will be activated to reduce the speed to close to zero and allow a controlled landing on the surface of Mars. A communication link between the EDM and the Orbiter will facilitate the real-time transmission of the most important data measured by the EDM. The ExoMars 2018 lander will be 2000kg at entry and includes a rover that will carry a drill and a suite of instruments dedicated to exobiology and geochemistry research. The parachute system challenges for this mission are greater given the increased system mass and unguided ballistic entry. ExoMars 2018 PAS trade studies led to a two stage system comprising a 15 m Disk Gap Band first stage parachute scaled from the ExoMars 2016 parachute and a 35 m ringslot second stage parachute. Both large parachutes are extracted by mortar deployed pilot chutes. The parachute system is initiated in the range Mach 1.9 Mach 2. Following deceleration to subsonic conditions on the first main parachute the second main parachute is deployed nominally at Mach 0.8. The front shield is separated 10s after second main parachute deployment. At the appropriate distance above the surface the lander separates from the backshell and parachute and will descend to the surface under the control of a liquid propulsion system. This presentation will discuss the design of both ExoMars parachute systems and their verification. The lander will touch down on Meridiani Planum, a relatively smooth, flat region on Mars, on 19 October This area interests scientists because it contains an ancient layer of hematite, an iron oxide that, on Earth, almost always forms in an environment containing liquid water. July 13-17, 2016 Laurel MD USA 6

7 Overview of the Parachute Decelerator System for the InSIGHT Mission Allen Witkowski Pioneer Aerospace Corporation 45 South Satellite Road South Windsor, CT David Buecher Lockheed Martin Space Systems South Wadsworth Blvd. Littleton, CO Devin Kipp Jet Propulsion Laboratory 4800 Oak Grove Drive El Segundo, CA Abstract The InSight spacecraft, consisting of a Cruise Stage and Lander, was planned to launch from Vandenberg AFB in March The Cruise Stage guides and powers the lander during the six month cruise to Mars. Once at Mars after a nominal six month transit, the Cruise Stage separates from the Lander, allowing the Lander to safely enter Mars, descend through the Martian atmosphere, and land on the surface. During the descent, the Lander will deploy the Disk- Gap Band (DGB) style parachute to decelerate it from supersonic to subsonic speeds. Once at the proper combination of altitude and velocity, the Lander is released from the Backshell (and parachute) and the Lander is then guided to, and lands on, the surface using several large thrusters. The entire Entry, Descent, and Landing (EDL) sequence relies heavily on the architecture and hardware developed, qualified, and used successfully during the Mars Scout Phoenix (PHX) mission, which launched in 2007 and landed in The InSight Parachute Decelerator System (PDS) is comprised of a Parachute Assembly and Mortar Assembly. The Parachute Assembly is based on the heritage 11.8m nominal diameter DGB Phoenix PDS. Both systems are nearly identical, with the only differences being in the textiles of the parachutes. The 625kg maximum entry mass of the InSight Lander is an increase from the 590kg maximum Phoenix entry mass. The increased entry mass, in combination with trajectory and atmospheric factors, led to an increase in the design parachute inflation load from 12,700lbf to 15,000lbf. To accommodate the increased peak parachute inflation load, the InSight parachute Suspension Line material has been changed from Kevlar (as used on Phoenix) to Technora and the Bridle Leg denier (strength) has been increased. This paper provides an overview of the InSight PDS program, with attention paid to the trades in design and development to apply current requirements for validation and verification to an existing or build to print design under cost cap and schedule constraints. Additional attention is paid to the alterations to canopy design in order to meet new mission strength performance and the lessons learned since its heritage PHX forebear. 1 July 13-17, 2016 Laurel MD USA 7

8 Decelerator Pack Rotation Discovery During Qualification / Acceptance Testing for the InSight Mission Allen Witkowski Pioneer Aerospace Corporation 45 South Satellite Road South Windsor, CT Al.Witkowski@ZodiacAerospace.com IPPW-13 Program Abstracts - Drag, Aerobraking and Aerocapture Session David Buecher Lockheed Martin Space Systems South Wadsworth Blvd. Littleton, CO david.buecher@lmco.com Devin Kipp Jet Propulsion Laboratory 4800 Oak Grove Drive El Segundo, CA devin.m.kipp@jpl.nasa.gov Abstract The Mars InSight program was scheduled for launch in the Spring of 2016, with an anticipated landing on Mars approximately six months later. As with previous Mars missions, the Entry, Descent, and Landing (EDL) architecture includes a parachute that will be deployed at supersonic speeds to provide the requisite deceleration and establish an appropriate descent timeline that supports all EDL functions. As described in an associated paper, 1 the InSight Parachute Decelerator System (PDS) is comprised of a Parachute Assembly and Mortar Assembly, and is nearly identical to the Mars Scout Phoenix (PHX) PDS, with the only noteworthy differences being in the textiles used to construct the parachutes. Additional similarities exist between the InSight and PHX missions, including comparable entry vehicles and PDS deployment conditions. Because of this, the InSight parachute should inflate rapidly in the low-density Martian atmosphere, in a manner that industry refers to as an Infinite Mass Inflation. Under this condition, vehicle deceleration is minimal during parachute inflation, and the peak parachute load occurs at or very near full inflation. Replicating this type of inflation process in the comparatively high density Earth atmosphere is among the most challenging aspects of any parachute structural qualification program. The Mars Exploration Rover (MER) and Mars Science Laboratory (MSL) parachute development and qualification programs utilized the NASA Ames National Full-Scale Aerodynamics Complex (NFAC) to better emulate an Infinite Mass Inflation on Earth. Unfortunately, the NFAC was closed shortly after MER, which relegated the PHX program to the more conventional aerial test method (which is prone to considerable vehicle deceleration and corresponding Non-Infinite Mass Inflation).. Fortunately, the NFAC was available for the InSight parachute structural qualification program, and advancements and lessons learned during the MSL program allowed for significant improvements in the sensors and cameras that were utilized. In fact, the data gathered during this NFAC test program, 2 was of sufficient quality and quantity to make informed comparisons between the InSight and PHX decelerator performance (despite the fact the PHX parachute system was tested under Non-Infinite Mass conditions). Immediately upon completion of the NFAC test program, a cursory review of high speed video from the final test revealed signs of deployment bag rotation and canopy twisting during the extraction process. Specifically, the last half of the canopy disk appeared twisted into a spiral. Follow-on review of all of the high speed video indicated that this phenomenon was present in each and every test, and the direction of rotation was the same as well. Parachute deployment is a critical function, and it is extremely fast in the low density, supersonic conditions associated with a flight mission. Therefore, twisting of the canopy could conceivably add risk to the flight mission, especially if it caused an abnormal inflated shape at elevated stress levels. This concern led to an investigation into the parachute twisting phenomenon. Of primary concern was understanding why and how the canopy developed twist, as well as what possible impact twist might have on the flight mission. Results from this investigation are presented along with implications for the InSight project and potential ramifications for future Mars missions. July 13-17, 2016 Laurel MD USA 8 1

9 PRELIMINARY STUDY ON MARS AEROCAPTURE AND LANDER MISSION USING DEEP SPACE MICRO SPACECRAFT AND MEMBRANE AEROSHELL. K. Suzuki 1, K. Yamada 2, R. Funase 3, H. Koizumi 4, I. Yoshikawa 4, K. Yoshioka 5, Y. Watanabe 3, D. Akita 6, O. Imamura 7 and S. Sugita 5, 1 Graduate School of Frontier Sciences, the University of Tokyo (5-1-5 Kashiwanoha, Kashiwa, Chiba , Japan, kjsuzuki@k.utokyo.ac.jp), 2 ISAS, JAXA (yamada.kazuhiko@jaxa.jp), 3 Graduate School of Engineering, the University of Tokyo, 4 Graduate School of Frontier Sciences, the University of Tokyo, 5 Graduate School of Science, the University of Tokyo, 6 Tokyo Institute of Technology, 7 Nihon University. Introduction: Nano, micro and small satellites are promising for not only commercial earth orbit utilization but also low-cost planetary exploration. In fact, the first interplanetary micro spacecraft PROCYON [1] (about 65 kg wet mass) was launched in 2014 and has been successfully demonstrating its performance of the micro ion thruster, the communication & navigation system, the bus system and so on for deep space missions. However, strict mass budget for the propulsion system does not allow us to make orbit insertion, though scientific results from observations from an insertion orbit are much more fruitful than those on a flyby trajectory. Use of the aerodynamic drag force for orbit insertion is known to be effective to save the spacecraft fuel. However, it requires highly accurate trajectory control, robustness against uncertainty in the properties of the planetary atmosphere, and the heat shield to protect the spacecraft from severe aerodynamic heating. Aerocapture has not been realized in the planetary exploration so far, because of its high risk. Aeroshells made from membrane materials have potential to solve the above problems. A vehicle with a large-area and small-mass membrane aeroshell enters the atmosphere at a small ballistic coefficient. It will receive sufficient deceleration even in low-density atmosphere of Mars, resulting in significant decrease in the aerodynamic heating. Due to low atmospheric density, ionization of the flow around the vehicle becomes weaker, and the communication blackout is expected to be avoided. Moreover, the membrane aeroshell can be compactly stored in the container, and can be deployed before the atmospheric entry by the inflatable system with high reliability. When we use the membrane aeroshell for the aerocapture, the success rate is expected be raised by appropriately controlling the time to separate the aeroshell from the spacecraft. The control error in the entry condition and the uncertainty in the atmospheric properties can be compensated for by such drag modulation technique. The advantages of the membrane aeroshell are not limited to a small spacecraft. However, difficulty to fabricate and handle the membrane aeroshell significantly increases with its size. Consequently, the combination of the membrane aeroshell and the PROCYON-class micro spacecraft seems promising to realize a small and lowcost Mars orbiter mission using aerocapture. The authors' research group has been developing the technologies in relation to the membrane aeroshell. In 2012, the suborbital reentry demonstration of the 1.2m-diameter inflatable flare-type membrane aeroshell (smaac) was successfully made from the altitude 150 km by using the JAXA's sounding rocket (S ) [2]. The 4-kg nano satellite EGG [3], which will be released from the International Space Station, is under preparation aming at the launch in 2016 to make the flight demonstration of the low-ballistic-coefficient orbital decay and the early stage of atmospheric entry. The objectives of the present study are 1) to create the mission scenario of the Mars exploration with a micro spacecraft and the membrane aeroshell aerocapture, and 2) to assess its feasibility. Fig. 1 Mission Scinario Mission Scinario: The mission scenario of the membrane aeroshell Mars exploration (hereinafter referred to as MAME ) is shown in Fig. 1. Before the atmospheric entry, the membrane aeroshell is deployed by the gas injection into the inflatable outer flame. At the same time, the inflatable nose heat shield is also deployed. The appropriate time for the aeroshell separation is autonomously determined by the on-board computer. After separation, the orbiter flies at a high ballistic coefficient without the aeroshell and goes out of the atmosphere. By firing the ion thruster, the periapsis altitude is raised to terminate the aerocapture. July 13-17, 2016 Laurel MD USA 9

10 The separated aeroshell continues to work as a landing probe. Due to extremely small ballistic coefficient after separation, it is quickly decelerated and is slowly descending in the atmosphere. Such slow descend is suitable for meteorological measurement and the observation of the surface by the cameras. Thanks to the absence of the blackout at the low-ballistic coefficient entry, the telecommunication between the orbiter and the lander is kept after the separation. At lower altitudes, the lander flies at the almost the same velocity as the wind. It will also work as a tracer for observation of the structure of the Martian wind. kw/m 2 by the method in [4]. At the entry path angle 7.5 degree, these are 160 Rm and 90.5kW/m 2. Such highly elliptic orbit after the aerodynamic pass provides sufficient time to raise the periapsis altitude by the ion thruster, and is suitable for the global observation of the upper atmosphere, the ionosphere, the solar wind interaction, the atmospheric escape and so on. At the entry path angle 9.0 degree, the peak heating rate becomes 351 kw/m 2, which is higher than the expected maximum service temperature of the membrane aeroshell under development. The increase in the maximum service temperature of the aeroshell results in raising the upper limit of the entry path angle. Concluding Remarks: In this paper, the concept of a micro spacecraft Mars exploration composed of an aerocapture orbiter and a lander using a membrane aeroshell MAME is presented. The low-ballisticcoefficient atmospheric entry of the membrane aeroshell vehicle has potential to widen the entry corridor for successful aerocapture by the drag modulation technique. The development of the membrane aeroshell is proceeding in a step-by-step manner as shown in Fig. 3. Fig. 2 Trajectory Analysis of MAME mission Trajectory Analysis: Typical results from the trajectory analysis on the MAME orbiter and the lander are shown in Fig. 2. We assume that the configuration of the MAME apacecraft at the atmospheric entry for the aerocapture is similar to that of our previous sounding rocket experiment [2]. The diameter of the nose heat shield is 0.66 m to put the orbiter with the same volume as the PROCYON satellite. The diameter of the aeroshell is 3.6 m. The total mass and the aeroshell mass are 80 kg and 10 kg, respectively. The ratio of the mass to the aeroshell front area is 7.9 kg/m 2 at the entry and 0.98 kg/m 2 after the separation. We use a simple criterion that the seroshell (lander) is separated from the orbiter at the time when the specific total energy E (E=V 2 /2-µ/r) becomes negative by receiving the aerodynamic drag force. The entry velocity is 5.5 km/s. Successful aerocapture and landing are obtained at entry flight path angles from 7.5 degree to 9.0 degree as shown in Fig. 2. At 8.0 degree, the obtained apoapsis altitude is 38 Rm (Rm is the radius of Mars) and the peak stagnation heating rate is estimated to be 134 Fig. 3 Development of Membrane Aeroshell References: [1] Funase R. (2015) 30th ISTS, Kobe, Japan, in World Space Highlight. [2] Yamada K. et al. (2015) J. Spacecraft and Rockets, 52(1), [3] Yamada K. et al. (2015) 30th ISTS, Kobe, Japan, 2015-f-61. [4] Tauber M. E. et al. (1990) J. Spacecraft and Rockets, 27(5), July 13-17, 2016 Laurel MD USA 10

11 DEVELOPMENT OF AN EARTH SMALLSAT FLIGHT TEST TO DEMONSTRATE VIABILITY OF MARS AEROCAPTURE. M. S. Werner 1, B.A. Woollard 2, A. Tadanki 3, R. D. Braun 4 and R. E. Lock 5, 1-4 Georgia Institute of Technology, Space Systems Design Laboratory, North Avenue NW, Atalnta, GA USA 1 mwerner9@gatech.edu, 5 Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA USA robert.e.lock@jpl.nasa.gov Introduction: Aerocapture has long been considered a compelling orbital maneuver that significantly reduces the cost of a wide variety of Mars orbital missions, as well as potential missions to Venus, Neptune, Titan, and return to Earth. Numerous conceptual studies have advanced the technical maturity of aerocapture to help enable its use on missions to different atmospheric worlds. Despite this technological readiness, the lack of an integrated aerocapture flight system demonstration has often been cited as rationale for not employing the technique on a flight mission. In order to facilitate a potential flight test, recent research has focused on the use of drag modulation-based techniques to greatly reduce the complexity of aerocapture systems. Drag modulation systems utilize changes in a vehicle s drag area during flight to effect control over the vehicle s ballistic coefficient, and therefore its final trajectory. Compared to traditional bank-to-steer lifting methods, these techniques enable use of extremely simple avionics algorithms, sensors and actuators, and eliminate the need for cg offset and an onboard propulsive reaction control system. Due to its simplicity, drag modulation-based aerocapture is a technique that can be readily tested via a smallsat, with results that can be applied to a variety of different planetary missions. This investigation is focused on the development of a conceptual smallsat mission that will demonstrate the feasibility of an aerocapture system. The spacecraft will be deployed as a secondary payload from a GTO; the conditions provided by the highenergy Earth-return trajectory from GTO will help minimize aerocapture targeting and post-maneuver delta-v requirements. Upon entering the atmosphere, the smallsat will employ hypersonic drag-modulation techniques to control the aerocapture maneuver: after enough energy has been lost via atmospheric drag, the spacecraft will jettison a drag shield, thereby modifying its ballistic coefficient and enabling capture into a relevant orbit after the atmospheric pass. Conceptual mission development has been focused on the design of the smallsat and the aerocapture device, as well as the modeling and selection of the spacecraft s atmospheric trajectory. The difference in ballistic coefficients before and after the drag shield is jettisoned will be the primary source of control authority over the aerocapture maneuver; as such, sizing both the main spacecraft body and the drag shield directly drives numerous other mission design aspects. Aerocapture system design is being accomplished through means of a comprehensive trade study, with focus will placed on control requirements, simplicity, and the applicability of the system to other missions. Three-degree-of-freedom numeric simulation can be used to model the spacecraft s trajectory, enabling quick analysis of peak heating and post-aerocapture orbits. Focus has also been placed on the scalability of results to different types of missions at Mars and other targets. This investigation will culminate in a baseline small satellite mission concept that is fully documented, including technical approach, management plan, cost estimation and risk assessment. A successful mission will show that drag modulation-based aerocapture can be used as an effective means of orbit insertion at Earth and other atmospheric worlds, with scalable applications to both large and small spacecraft. July 13-17, 2016 Laurel MD USA 11

12 IMPROVING BLUNT BODY AERODYNAMIC PERFORMANCE THROUGH DEPLOYABLE TABS. A. M. Korzun 1, 1 NASA Langley Research Center, M/S 486, Hampton, VA 23681, ashley.m.korzun@nasa.gov. Introduction: Achieving precision landing, accessing lander sites at higher surface elevations, and increasing payload mass to the surface of Mars require the ability to generate aerodynamic lift [1,2]. The Viking landers (Viking I and Viking II) and Mars Science Laboratory (MSL) are the only vehicles flown to date at Mars that were designed to fly with lift. All three vehicles provided aerodynamic lift with an axisymmetric blunt body through a radial offset of the vehicle s center of gravity. The Viking landers achieved this radial CG offset through a combination of the stowed lander s position within the aeroshell and ballast mass [3,4]. MSL used six tungsten entry ballast masses, totaling 174 kg, to achieve the required trim angle of attack during entry, with all six masses jettisoned prior to parachute deployment to re-trim the vehicle to a near-zero angle of attack. An additional kg of ballast were required to balance the MSL cruise stage, for a total ballast mass of kg [5]. For comparison, the mass of each Mars Exploration Rover was 174 kg. Recent mechanical and thermal design to integrate an appropriately sized tab into the MSL entry vehicle indicated the possibility of achieving the same trim performance with at deployable tab at 1/10 th of the mass that would have been required as ballast [6]. Constraints on the entry, descent, and landing (EDL) system could be reduced if similar or improved aerodynamic performance can be achieved with a less massive system. In addition to robotic scale missions, increased aerodynamic performance is of interest to extend the applicability of axisymmetric blunt body configurations to much larger payloads, such as those required for the human exploration of the surface of Mars. Trim tabs are aerodynamic control surfaces that can allow an entry vehicle to achieve aerodynamic performance requirements without needing an offset center of gravity. These deployable surfaces can provide the ability to modulate L/D or even ballistic coefficient, and a combination of tabs can provide increased control authority in both crossrange and downrange. An example of an aeroshell with a single trim tab is shown in Fig. 1. The number of tabs, location of tabs radially about the vehicle, tab area, tab cant angle, tab aspect ratio are all parameters that can be traded to achieve a desired performance metric or range of performance metrics. This paper reviews prior investigations into the use of trim tabs for improving the aerodynamic performance of blunt bodies during atmospheric entry and provides an overview of present interest in their application. The characteristics of tabs to achieve a specified range performance are also discussed. Figure 1. Example of a mechanically deployed rigid trim tab [6]. History: Trim tabs or deployable flaps have been of interest to the EDL community since the early 1960s [7-10]. While such control surfaces were never incorporated into final designs, a number of tests were done using various sizes and arrangements of flaps, and this early work provided guidance to more recent investigations with regard to where the tab(s) should be placed on the vehicle and how large the tab(s) needed to be to achieve a particular trim condition. Since this time, additional experimental work and systems analysis have been completed, with particular interest in potential applications to the Mars Smart Lander, which later became the successful Mars Science Laboratory mission. The full version of this paper will discuss the available literature, ranging from the 1960s through the present. Uses: Deployable tabs trim the vehicle to a nonzero angle, either an angle of attack or sideslip, without requiring an offset of the vehicle s center of gravity. The tab (or tabs), when deployed, provide additional surface area with a significant moment arm, disrupting the symmetry and altering the trim characteristics of the vehicle. A single tab provides uni-directional control, while two or more tabs can provide directional control in both pitch and yaw. As the tabs are integrated with the shoulder of the vehicle, interactions with the vehicle s wake are avoided, and trim tabs may provide more responsiveness in crossrange than a traditional propulsive reaction control system and better performance across a wider range of Mach numbers during entry and descent. The full paper will present July 13-17, 2016 Laurel MD USA 12

13 specifics on the aerodynamic performance achievable with deployable tabs. References: [1] Lockwood M. K. et al. (2006) Journal of Spacecraft and Rockets, Vol. 43, No. 2, [2] Korzun A. M. et al. (2013) AIAA [3] Viking 79 Rover Study Volume 1: Summary Report (1974) NASA CR [4] Edquist K. T. (2006) AIAA [5] Schoenenberger, M. et al. (2009) AIAA [6] Ivanov, M. C. et al. (2011) NASA/TM [7] Sammonds R. I. and Dickey R. R. (1961) NASA TM X-579. [8] Tendeland T. and Pearson, Jr. B. D. (1962) NASA TM X-660. [9] Neal, Jr. L. (1963) NASA TM X-816. [10] Armstrong, W. O. (1961) NASA TM X-469. July 13-17, 2016 Laurel MD USA 13

14 USE OF PLASMA ACTUATOR TO IMPROVE REENTRY MISSIONS. S. Coumar 1 and V. Lago 2, 1,2 ICARE, CNRS, UPR 3021, 1C Avenue de la Recherche Scientifique, 45071, Orléans Cedex 2, France. 1 sandra.coumar@cnrs-orleans.fr 2 viviana.lago@cnrs-orleans.fr Introduction: During atmospheric entry missions, three requirements must be balanced: deceleration, heat management and accuracy of the localization and velocity when landing. These requirements strongly depend on the drag force which is the dominant force over all others including gravity and lift. On this purpose, it has been demonstrated that plasma actuators offer the possibility to modify the shock wave shape around the vehicle and thus drag coefficient. The goal of the presentation is to track the work done with a plasma actuator in a supersonic flow since 2005 in the ICARE laboratory. Methodology: This study is mainly experimental and is carried out with the super/hypersonic wind tunnel MARHy located at the ICARE laboratory in Orléans (France). It is focused on an aluminium electrode mounted on a beveled flat plate. The model is exposed to a rarefied supersonic flow, inducing the creation of a glow discharge when firing the electrode. The interaction between the ionization and the flow is then studied through several measuring methods as flow visualisation, thermal analysis, Langmuir probes,... Drawing of the experimental setup Findings: It is shown that when using the plasma actuator, the shock angle occurring when the supersonic flow hit the model is increasing. For example, when generating a Mach 2 flow, for a discharge current of 58 ma, the shock angle would go from 36 (when the discharge is off) to 42. In order to explain the phenomenon, different hypothesis were proposed. First, it was thought that this increase of the shock wave angle was the cause of an icrease of the gas temperature, i.e the volumetric heating. It was shown through spectroscopy measurements that this volumetric heating is accounting only for a little in the shock wave increase[1]. Then, the surface heating, due to the Joule effect was studied with the conception of an heating element. The result was that, although this type of heating causes an increase of the shock wave angle, it is not enough to explain the values obtained with the plasma actuator [2]. Therefore, it is clear that another phenomenon, more than heating arises. Observation is made that the ionization rate is lower with the ionization of the gas and that the ionization itself has an important role in the shock wave modification. Image of the flow field modified by the plasma actuator (V HV = kv and I HV = 74 ma) Application: Drag and lift forces have been estimated through calculations. For a Mach 2 flow at a static pressure of 8 Pa, the drag force is increasing as expected and the drag coefficient is modified by +13% for the plasma actuation, in comparison to only +5% with a pure surface heating [3]. This result can be linked to the heat flux over the spaceship. If taking for model a spaceship in the mold of the space shuttle Columbia, an increase of 13% of the drag coefficient would lead to a decrease of the vehicle speed of about 7% and of about 26% of the heat flow. References: [1] Menier E. (2007), PhD Thesis, University of Orléans. [2] Joussot R. et al. (2015) Exp. in Fluids, vol. 56, pp [3] Coumar S. et al. (2015) AIAA 20 th International Space Planes and Hypersonic Systems and Technologies Conference, Glasgow, UK. July 13-17, 2016 Laurel MD USA 14