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Development of an Earth Smallsat Flight Test to Demonstrate Viability of Mars Aerocapture
... Recent work has highlighted the opportunity to leverage the significant development in CubeSat and SmallSat flight system hardware for an extremely cost-effective technology demonstration of the drag modulation system in Earth orbit [29]. A small (less than 100 kilogram) aerocapture spacecraft would use either a secondary launch opportunity or a small launch vehicle to a highly elliptical orbit to perform a single atmospheric pass at relevant conditions to planetary entry. ...
Aerocapture technology provides a far-reaching capability for missions at destinations across the solar system, from SmallSat to Flagship class. It will enable a new class of SmallSat orbiters, allow high priority missions to carry more science payload to their destination, and open a new rapid transportation pathway to the outer solar system. Significant technology development and demonstration in the past decade has prepared high heritage aeroshell shapes to be ready to support the infusion of aerocapture into missions, particularly at the Ice Giants. In addition, research and focus on drag modulation control schemes has highlighted that this aerocapture trajectory control system can enable a new generation of small spacecraft science orbiters as secondary payloads on larger mission’s launches under programs such as SIMPLEx, leading to increased science return
at a fraction of the cost.
... Recent work has highlighted the opportunity to leverage the significant development in CubeSat and SmallSat flight system hardware for an extremely cost-effective technology demonstration of the drag modulation system in Earth orbit [29]. A small (less than 100 kilogram) aerocapture spacecraft would use either a secondary launch opportunity or a small launch vehicle to a highly elliptical orbit to perform a single atmospheric pass at relevant conditions to planetary entry. ...
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... Despite this simplicity, previous analyses have shown that discrete drag modulation systems are capable of achieving accuracy competitive with that of lift-modulation systems. 1,4,5 Additionally, drag-modulation systems typically have low ballistic coefficients, resulting in a more benign aerothermal environment than that experienced by lifting aerocapture concepts, enabling drag modulation systems to utilize less expensive and lighter weight thermal protection system materials. ...
Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of a small satellite aerocapture capability would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite aerocapture system is assessed for use at Mars across a range of small satellite payloads, approach trajectories, and destinations in the Mars system. The aerocapture system uses drag modulation for trajectory control to ensure successful orbit insertion. Analyses include assessment of the sensitivity of the entry corridor size to the ballistic-coefficient ratio, the effectiveness of real-time aerocapture guidance and control algorithms, aerocapture system-level impacts of different target orbits, and development of requirements and recommendations for the development of a small satellite aerocapture system. Results indicate that a discrete drag-modulation aerocapture system may provide an orbit-insertion capability for small satellites with modest propulsion requirements.
... One possible way to execute the single-stage discrete-event maneuver is to enter the atmosphere in a low ballistic coefficient configuration, with a large drag skirt deployed, and then to transition to high ballistic coefficient by jettisoning or folding the drag skirt. Such a single-stage discrete-event architecture was previously studied within the context of an Earth SmallSat Flight Test of Aerocapture [6]. Additional information on the fundamentals of drag modulation aerocapture can be found in Reference 5. ...
... Putnam and Braun proposed the concept of the drag modulation flight control system and analyzed the guidance performance using NPC [7]. Michael et al. has developed an Earth smallsat flight test for a Mars aerocapture [18]. Han et al. has proposed an optimal ballistic coefficient control [19]. ...
Aerocapture can significantly reduce the velocity increment required for a planetary orbital mission and reduce the amount of propellant needed. And it may be one of the key technologies necessary for large-scale space exploration missions in the future. In this paper, the analytical solution of aerocapture based on the piecewise variable ballistic coefficient is studied around the exploration of Mars. An aerocapture analytical predictive guidance algorithm for single ballistic coefficient switching is proposed. The terminal velocity after the ballistic coefficient switching can be obtained by analytical calculation in real time. The adaptive control of the switching time of the ballistic coefficient is realized. The simulation results show that the guidance algorithm is accurate and robust, which can effectively overcome the influence of atmospheric density error, aerodynamic parameter error, and initial state uncertainty.
Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of an independent orbit insertion capability for small satellites, using aerocapture, would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite drag-modulation aerocapture system is assessed for use at Mars across a range of approach trajectories and destinations in the Mars system. Analyses include assessment of the sensitivity of the entry corridor size over different atmospheric conditions, a comparison of velocity-trigger and numerical predictor-corrector guidance schemes for drag modulation, and aerocapture flight performance assessment via Monte Carlo techniques. A special focus is placed on four baseline missions: a low-altitude Mars mapping orbit, Phobos and Deimos flyby/rendezvous, and areosynchronous orbit. Results indicate that aerocapture may decrease the orbit insertion system mass fraction by 30% or more with respect to fully propulsive options.
Discrete-event drag-modulation trajectory control is assessed for planetary entry using the closed-form Allen-Eggers solution to the equations of motion. A control authority metric for drag-modulation trajectory control systems is derived. Closed-form analytical relationships are developed to assess range divert capability and to identify jettison condition constraints for limiting peak acceleration and peak heat rate. Closed-form relationships are also developed for drag-modulation systems with an arbitrary number of stages.
The selection of the unique aeroshell shape for the Mars Microprobes is discussed. A description of its aerodynamics in hypersonic rarefied, hypersonic continuum, supersonic and transonic flow regimes is then presented. This description is based on Direct Simulation Monte Carlo analyses in the rarefied-flow regime, thermochemical nonequilibrium Computational Fluid Dynamics in the hypersonic regime, existing wind tunnel data in the supersonic and transonic regime, additional computational work in the transonic regime, and finally, ballistic range data. The aeroshell is shown to possess the correct combination of aerodynamic stability and drag to convert the probe's initial tumbling attitude and high velocity at atmospheric-interface into the desired surface-impact orientation and velocity.
Successful return of interstellar dust and cometary material by the Stardust Sample Return Capsule requires an accurate description of the Earth entry vehicle's aerodynamics. This description must span the hypersonic-rarefied, hypersonic-continuum, supersonic, transonic, and subsonic flow regimes. Data from numerous sources are compiled to accomplish this objective. These include Direct Simulation Monte Carlo analyses, thermochemical nonequilibrium computational fluid dynamics, transonic computational fluid dynamics, existing wind tunnel data, and new wind tunnel data. Four observations are highlighted: 1) a static instability is revealed in the free-molecular and early transitional-flow regime due to aft location of the vehicle s center-of-gravity, 2) the aerodynamics across the hypersonic regime are compared with the Newtonian flow approximation and a correlation between the accuracy of the Newtonian flow assumption and the sonic line position is noted, 3) the primary effect of shape change due to ablation is shown to be a reduction in drag, and 4) a subsonic dynamic instability is revealed which will necessitate either a change in the vehicle s center-of-gravity location or the use of a stabilizing drogue parachute.
A software tool for the prediction of the aero -thermodynamic environments of conceptual aerospace configurations is presented. The vehic le geometry is defined using unstructured, triangulated surface meshes. For subsonic Mach numbers a fast, unstructured, multi -pole panel code is coupled with a streamline tracing formulation to define the viscous surface solution. For supersonic and hypers onic Mach numbers, various independent panel methods are coupled with the streamline tracing formulation, an attachment line detection method , and stagnation -attachment line heating models to define the viscous aero -thermal environment.
Using an inflatable ballute system for aerocapture at planets and moons with atmospheres has the potential to provide significant performance benefits compared not only to traditional all propulsive capture, but also to aeroshell based aerocapture technologies. This paper discusses the characteristics of entry trajectories for ballute aerocapture at Neptune. These trajectories are the first steps in a larger systems analysis effort that is underway to characterize and optimize the performance of a ballute aerocapture system for future missions not only at Neptune, but also the other bodies with atmospheres.
Drag-modulation flight control may provide a simple method for controlling energy during aerocapture. Several drag-modulation flight-control system options are discussed and evaluated, including single-stage jettison, two-stage jettison, and continuously variable drag-modulation systems. Performance is assessed using numeric simulation with real-time guidance and targeting algorithms. Monte Carlo simulation is used to evaluate system robustness to expected day-of-flight uncertainties. Results indicate that drag-modulation flight control is an attractive option for aerocapture systems at Mars, where low peak heat rates enable the use of lightweight inflatable drag areas. Aerocapture using drag modulation at Titan is found to require large drag areas to limit peak heat rates to nonablative thermal-protection system limits or advanced lightweight ablators. The large gravity well and high peak heat rates experienced during aerocapture at Venus make drag-modulation flight control unattractive when combined with a nonablative thermal-protection system. Significantly larger drag areas or advances in fabric-based material thermal properties are required to improve feasibility at Venus.
Hypersonic deployable aerodynamic devices, both rigid and inflatable, have the potential to enable a broad spectrum of next-generation aeroassist missions by mitigating shape and size constraints on aeroassist vehicles and providing an in-flight reconfiguration capability. Such a capability provides new options for flight control during atmospheric flight, such as drag modulation. Drag modulation is an attractive flight control option for future aerocapture missions because it requires only minimal additional system complexity for vehicles with deployable aerodynamic devices, in contrast to more conventional lift-modulation steering methods. This study expands upon previous aerocapture drag modulation studies by extending the analysis of single-event jettison systems to Earth and Mars. A single-event jettison guidance algorithm was developed and used to evaluate the feasibility of real-time targeting of apoapsis altitude during flight. Results indicate that sufficiently large ballistic coefficient ratios provide adequate aerodynamic and guided corridors for future aerocapture missions. While the preliminary guidance algorithm demonstrates only modest insertion accuracy, this level of accuracy may be tolerable for certain missions.
A Neptune Aerocapture Systems Analysis is completed to determine the feasibility, bene- fit and risk of an aeroshell aerocapture system for Neptune and to identify technology gaps and technology performance goals. The high fidelity systems analysis is completed by a five center NASA team and includes the following disciplines and analyses: science; mission de- sign; aeroshell configuration screening and definition; interplanetary navigation analyses; atmosphere modeling; computational fluid dynamics for aerodynamic performance and da- tabase definition; initial stability analyses; guidance development; atmospheric flight simula- tion; computational fluid dynamics and radiation analyses for aeroheating environment definition; thermal protection system design, concepts and sizing; mass properties; struc- tures; spacecraft design and packaging; and mass sensitivities. Results show that aerocapture can deliver 1.4 times more mass to Neptune orbit than an all-propulsive system for the same launch vehicle. In addition aerocapture results in a 3-4 year reduction in trip time compared to all-propulsive systems. Aerocapture is feasible and performance is adequate for the Neptune aerocapture mission. Monte Carlo simulation re- sults show 100% successful capture for all cases including conservative assumptions on at- mosphere and navigation. Enabling technologies for this mission include TPS manufactur- ing; and aerothermodynamic methods and validation for determining coupled 3-D convec- tion, radiation and ablation aeroheating rates and loads, and the effects on surface recession.
Calculations have been performed to quantify the cost and delivered mass advantages of aerocapture at all destinations in the Solar System with significant atmospheres. A total of eleven representative missions were defined for the eight possible destinations and complete launch-to-orbit insertion architectures constructed. Direct comparisons were made between aerocapture and competing orbit insertion techniques based on state-of-the-art and advanced chemical propulsion, solar electric propulsion, and aerobraking. The results show that three of the missions cannot be done without aerocapture: Neptune elliptical orbits, Saturn circular orbits, and Jupiter circular orbits. Aerocapture was found to substantially reduce the cost per unit mass delivered into orbit for five other missions based on a heavy launch vehicle: Venus circular orbits (55% reduction in $/kg costs), Venus elliptical orbits (43% reduction); Mars circular orbits (13% reduction), Titan circular orbits (75% reduction), and Uranus circular orbits (69% reduction). These results were found to be relatively insensitive to 30% increases in both the estimated aerocapture system mass
The paper is concerned with the aeroassisted orbit transfer, i. e. , using aerodynamic forces to produce an orbital plane change with an expenditure of energy significantly smaller than that associated with an extra-atmospheric propulsive maneuver. A review is made of some of the numerous studies of aeroassisted orbit transfer carried out so far for a wide variety of Earth-orbital and planetary missions. Almost any mission that involves changes in orbital altitude and/or inclination in the vicinity of an atmosphere-bearing planet is a candidate for aeroassist. For ease of discussion these many possible mission applications may be categorized as synergetic plane change, planetary mission applications, and orbital transfer vehicle applications. However, aeroassist has seldom been employed in actual missions. The various vehicle concepts are examined in greater detail and an attempt is made to identify the technology barriers that have, thus far, prevented aeroassisted orbit transfer from achieving practical implementation.