David Kinney’s research while affiliated with Ames Research Center and other places

View Video Presentation: https://doi.org/10.2514/6.2022-0417.vid As interest in non-traditional entry vehicles continues to grow, the need for a Thermal Protection System (TPS) analysis approach that can account for entry solutions with changing geometry and complicated flow dynamics becomes invaluable. The NASA Space Technology Mission Directorate (STMD) Pterodactyl project aims to accomplish this through examination of a flap controlled Adaptable, Deployable Entry Placement Technology (ADEPT)-style Deployable Entry Vehicle (DEV). This paper details an improved methodology for modeling the aerothermodynamic environment and initial TPS design for a flap control system integrated with a symmetric DEV. Improvements include i) the addition of an anchoring process that uses an increased fidelity aerodynamic solution to anchor the aerothermal environment predictions and ii) increased surface resolution, for the 1D heat transfer analysis, to isolate the hottest area on the flap that will require the thickest TPS. It was found that the anchoring process provided an improved aerothermal environment prediction. Additionally, this analysis demonstrated that there is a need for increased surface resolution in the 1D thermal analysis since the predicted location of the hottest point was significantly different than the location identified using very coarse surface resolution.

A method and associated system for multi-disciplinary optimization of various parameters associated with a space vehicle that experiences aerocapture and atmospheric entry in a specified atmosphere. In one embodiment, simultaneous maximization of a ratio of landed payload to vehicle atmospheric entry mass, maximization of fluid flow distance before flow separation from vehicle, and minimization of heat transfer to the vehicle are performed with respect to vehicle surface geometric parameters, and aerostructure and aerothermal vehicle response for the vehicle moving along a specified trajectory. A Pareto Optimal set of superior performance parameters is identified.

The state-of-the-art in vehicle design decouples trajectory generation from shape optimization, which may result in an aeroshell design that does not meet in-flight requirements. The integration of a guidance algorithm into the design process can provide a real-time, rapid trajectory generation technique to improve vehicle design solutions. This work quantifies the performance of a reference tracking guidance algorithm, an Apollo Derived Guidance (ADG), for five different geometries using a single same reference trajectory. The guided trajectories are compared to the trajectories determined in a vehicle optimization study for a Mars Entry. The aerodynamic dispersions ranged from +/-1% to +/-17% and a single extreme case applies an aerodynamic dispersion of approximately 80% less than the baseline geometry. This study revealed that the generation of flight feasible trajectories is only as good as the robustness of the guidance algorithm. The ADG, as expected, was able to guide the vehicle into the separation box at the target location for dispersions up to 17%, but failed for the 80% dispersion case. Finally, the results revealed that including flight feasible trajectories for a set of dispersed geometries has the potential to save up to 430 kg of mass.

The present paper describes an innovative, semi-rigid, mechanically deployable hypersonic decelerator system for human missions to Mars. The approach taken in the present work builds upon previous architecture studies performed at NASA, and uses those findings as the foundation to perform analysis and trade studies. The broad objectives of the present work are: (i) to assess the viability of the concept for a heavy mass (landed mass ≈40 mT) Mars mission through system architecture studies; (ii) to contrast it with system studies previously performed by NASA; and (iii) to make the case for a Transformable Entry System Technology. The mechanically deployable concept at the heart of the proposed transformable architecture is akin to an umbrella, which in a stowed configuration meets launch requirements by conforming to the payload envelope in the launch shroud, and when deployed in earth orbit forms a large aerosurface designed to provide the necessary aerodynamic forces upon entry into the Martian atmosphere. The aerosurface is a thin skin draped over high-strength ribs; the thin skin or fabric with flexible material serves as the thermal protection system, and the ribs serve as the structure. A four-bar linkage mechanism allows for a reorientation of the aerosurface during aerocapture or during the entry and descent phases of atmospheric flight, thus providing a capability to navigate and control the vehicle and make possible precision landing. The actuators and mechanisms that are used to deploy the aerosurface are multi-functional in that they also allow for reorienting the aerosurface during entry, descent and landing phases. The structure together with the TPS is actively inverted in a controlled manner soon after the vehicle slows down from ignition of the retro-propulsion engines, i.e., the aerosurface is inverted like an umbrella to form a landing system. The inverted aerosurface also serves as a shield for the payload and surrounding critical components by providing debris protection. At landing, the structural spokes of the umbrella provide landing impact attenuation, thus assisting in soft and safe landing on uneven terrain.