Hypersonic flow analysis is performed on an inflatable aerocapture device called a "Ballute" for Titan's Mission. An existing unstructured Cartesian grid methodology is used as a starting point by taking advantage of its ability to automatically generate grids over any deformed shape of the flexible ballute. The major effort for this thesis work is focused on advancing the existing unstructured Cartesian grid methodology. This includes implementing thermochemical nonequilibrium capability and porting it to a parallel computing environment using a Space-Filling-Curve (SFC) based domain decomposition technique. The implemented two temperature thermochemical nonequilibrium solver governs the finite rate chemical reactions and vibrational relaxation in the high temperature regimes of hypersonic flow. In order to avoid the stiffness problem in the explicit chemical solver, a point implicit method is adopted to calculate the chemical reaction source term. The AUSMPW+ scheme with MUSCL data reconstruction is adopted as the numerical scheme to avoid non-physical oscillations and the carbuncle phenomenon. The results for five species air model and for thirteen species N2-CH4-Ar model to simulate Titan entry are included for verification against DPLR (NASA Ames' structured grid hypersonic flow solver). The efficient parallel computation of any unstructured grid flow solver requires an adequate grid decomposition strategy because of its complex spatial data structure. The difficulties of even and block-contiguous partitioning in frequently adapting unstructured Cartesian grids are overcome by implementing the 3D Hilbert SFC. Grids constructed by the SFC for parallel environment promise short inter-CPU communication time while maintaining perfect load balancing between CPUs. The load imbalance due to the local solution adaption is simply apportioned by re-segmenting the curve into even pieces. The detailed structure of the 3D Hilbert SFC and parallel computing efficiency results based on this grid partition method are also presented. Finally, a structural dynamics tool (LS-DYNA) is loosely coupled with the present parallel thermochemical nonequilibrium flow solver to obtain the deformed surface definition of the ballute.

Aerocapture is a method for spacecraft orbital insertion that is currently being assessed for use in interplanetary missions. This method would use a low periapsis hyperbolic entry orbit to induce drag allowing the spacecraft to slow down without the use of a propulsion system. This is accomplished by using a ballute (balloon parachute), which is released after the appropriate change in velocity necessary to achieve the desired planetary orbit. Once released, the ballute could deploy a secondary mission vehicle. A MATLAB simulation was run to understand the environment a secondary payload would undergo, such as heating and deceleration, as well as to study the buoyancy due to the ballute. The stability of the spacecraft during entry is also discussed. The results showed that if the ballute can survive the aerocapture maneuver then it will be able to survive entry with a secondary payload. The deceleration from the separation of the primary and secondary payload will be large but it can be overcome. The stability of the vehicle is dependent on the location of the center of gravity. Buoyancy at Mars has little effect due to the low density of the atmosphere; at higher density atmospheres buoyancy does play a role in the payload descent. Results of the analysis show that a successful landing of a ballute with a secondary payload is possible.

Ballute hypersonic aerodynamic decelerators have been considered for aerocapture since the early 1980's. Recent technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. Several technology hurdles have been identified, including the effects of coupled fluid structure interaction on ballute performance and survivability. To date, no aeroelastic solutions of thin-film ballutes in an environment relevant to aerocapture have been published. In this investigation, an aeroelastic solution methodology is presented along with the analysis codes selected for each discipline. Variable-fidelity aerodynamic tools are used due to the long run times for computational fluid dynamics or direct simulation Monte Carlo analyses. The improved serial staggered method is used to couple the disciplinary analyses in a time-accurate manner, and direct node-matching is used for data transfer. In addition, an engineering approximation has been developed as an addition to modified Newtonian analysis to include the first-order effects of damping due to the fluid, providing a rapid dynamic aeroelastic analysis suitable for conceptual design. Static aeroelastic solutions of a clamped ballute on a Titan aerocapture trajectory are presented using non-linear analysis in a representative environment on a flexible structure. Grid convergence is demonstrated for both structural and aerodynamic models used in this analysis. Static deformed shape, drag and stress level are predicted at multiple points along the representative Titan aerocapture trajectory. Results are presented for verification and validation cases of the structural dynamics and simplified aerodynamics tools. Solutions match experiment and other validated codes well. Contributions of this research include the development of a tool for aeroelastic analysis of thin-film ballutes which is used to compute the first high-fidelity aeroelastic solutions of thin-film ballutes using inviscid perfect-gas aerodynamics. Additionally, an aerodynamics tool that implements an engineering estimate of hypersonic aerodynamics with a moving boundary condition is developed and used to determine the flutter point of a thin-film ballute on a Titan aerocapture trajectory.

Since the 1970's, NASA has relied on the use of rigid aeroshells and supersonic parachutes to enable robotic mission to Mars. These technologies are constrained by size and deployment condition limitations that limit the payload they can deliver to the surface of Mars. One candidate technology envisioned to replace the supersonic parachute is the supersonic inflatable aerodynamic decelerator (IAD). This dissertation presents an overview of work performed in maturing a particular type of IAD, the tension cone. The tension cone concept consists of a flexible shell of revolution that is shaped so as to remain under tension and resist deformation. Systems analyses that evaluated trajectory impacts of a supersonic IAD demonstrated several key advantages including increases in delivered payload capability of over 40%, significant gains in landing site surface elevation, and the ability to accommodate growth in the entry mass of a spacecraft. A series of supersonic wind tunnel tests conducted at the NASA Glenn and Langley Research Centers tested both rigid and flexible tension cone models. Testing of rigid force and moment models and pressure models demonstrated the new design to have favorable performance including drag coefficients between 1.4 and 1.5 and static stability at angles of attack from 0º to 20º. A separate round of tests conducted on flexible tension cone models showed the system to be free of aeroelastic instability. Deployment tests conducted on an inflatable model demonstrated rapid, stable inflation in a supersonic environment. Structural modifications incorporated on the models were seen to reduce inflation pressure requirements by a factor of nearly two. Through this test program, this new tension cone IAD design was shown to be a credible option for a future flight system. Validation of CFD analyses for predicting aerodynamic IAD performance was also completed and the results are presented. Inviscid CFD analyses are seen to provide drag predictions accurate to within 6%. Viscous analyses performed show excellent agreement with measured pressure distributions and flow field characteristics. Comparisons between laminar and turbulent solutions indicate the likelihood of a turbulent boundary layer at high supersonic Mach numbers and large angles of attack.

The current limits of supersonic parachute technology may constrain the ability to safely land future robotic assets on the surface of Mars. This constraint has led to a renewed interest in supersonic inflatable aerodynamic decelerator (IAD) technology, which offers performance advantages over the DGB parachute. Two supersonic IAD designs of interest include the isotensoid and tension cone, named for their respective formative structural theories. Although these concepts have been the subject of various tests and analyses in the 1960s, 1970s, and 2000s, significant work remains to advance supersonic IADs to a technology readiness level that will enable their use on future flight missions. In particular, a review of the literature revealed a deficiency in adequate aerodynamic and aeroelastic data for these two IAD configurations at transonic and subsonic speeds. The first portion of this research amended this deficiency by testing flexible IAD articles at relevant transonic and subsonic conditions. The data obtained from these tests showed that the tension cone has superior drag performance with respect to the isotensoid, but that the isotensoid may demonstrate more favorable aeroelastic qualities than the tension cone. Additionally, despite the best efforts in test article design, there remains ambiguity regarding the accuracy of the observed subscale behavior for flight scale IADs. Due to the expense and complexity of large-scale testing, computational fluid-structure interaction (FSI) analyses will play an increasingly significant role in qualifying flight scale IADs for mission readiness. The second portion of this research involved the verification and validation of finite element analysis (FEA) and computational fluid dynamic (CFD) codes for use within an FSI framework. These verification and validation exercises lend credence to subsequent coupled FSI analyses involving more complex geometries and models. The third portion of this research used this FSI framework to predict the static aeroelastic response of a tension cone IAD in supersonic flow. Computational models were constructed to mimic the wind tunnel test articles and flow conditions. Converged FSI responses computed for the tension cone agreed reasonably well with wind tunnel data when orthotropic material models were used and indicated that current material models may require unrealistic input parameters in order to recover realistic deformations. These FSI analyses are among the first results published that present an extensive comparison between FSI computational models and wind tunnel data for a supersonic IAD.