Inflatable Aerodynamic Decelerator Research Articles

Aerothermodynamic analysis for deformed membrane of inflatable aeroshell in orbital reentry mission

An inflatable aerodynamic decelerator with a membrane aeroshell is a promising key technology in the reentry, descent, and landing phases of future space transportation. The membrane aeroshell is generally deformed by the in-flight aerodynamic force; however, the effects of the deformation on the aerodynamic heating are unclear. Here, we investigated aerodynamic heating for an inflatable reentry vehicle, Titans, in the hypersonic regime using flow field simulation coupled with structural analysis. Thermochemical nonequilibrium flows around the Titans with a deformed membrane aeroshell were reproduced numerically for an angle of attack (AoA) values between 0° and 40°. The maximum displacements of the membrane aeroshell by deformation at the AoAs of 0° and 40° were 6.7% and 6.6% of the diameter of the Titans, respectively. The difference in heat fluxes between the deformed and rigid shapes was a remarkable 188.8% for a 0° AoA owing to the considerable changes in the front shock wave shape. Meanwhile, it was indicated that membrane deformation at an AoA of 40° insignificantly affected the peak heat flux value on the inflatable torus because the considerable change in the shock wave shape observed for the case of 0° AoA did not occur. It was found that local wrinkles on the membrane aeroshell were formed by deformation, thus causing the heat flux to increase owing to an increase in local temperature gradient on the surface.

Modeling inflatable fabric with undevelopable surfaces by motion folding method

At present, most space inflatable structures are composed of flexible inflatable fabrics with complex undevelopable surfaces. It is difficult to establish a multi-dimensional folding model for this type of structure. To solve this key technical problem, the motion folding method is proposed in this study. First, a finite element model with an original three-dimensional surface was flattened with a fluid structure interaction algorithm. Second, the flattened surface was folded based on the prescribed motion of the node groups, and the final folding model was obtained. The fold modeling process of this methodology was consistent with the actual folding processes. Because the mapping relationship between the original finite element model and the final folding model was unchanged, the initial stress was used to modify the model errors during folding process of motion folding method. The folding model of an inflatable aerodynamic decelerator, which could not be established using existing folding methods, was established by using motion folding method. The folding model of the inflatable aerodynamic decelerator showed that the motion folding method could achieve multi-dimensional folding and a high spatial compression rate. The stability and regularity of the inflatable aerodynamic decelerator numerical inflation process and the consistency of the inflated and design shapes indicated the reliability, applicability, and feasibility of the motion folding method. The study results could provide a reference for modeling complex inflatable fabrics and promote the numerical study of inflatable fabrics.

Assessment of laminar-turbulent transition models for Hypersonic Inflatable Aerodynamic Decelerator aeroshell in convection heat transfer

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) has shown its great potential for future planetary explorations. However, the HIAD surface deflections could both promote boundary-layer transition early and augment heating levels sharply, which poses challenges for survivability of a Thermal Protection System (TPS). The goal of this work is to assess different transition models for the prediction of hypersonic transitional flows over scalloping deformed HIAD surface and to seek the critical factor for their capabilities to provide a reference for the application and advancement. Three representative transition models are considered: γ-Reθ, k-ω-γ and kT-kL-ω. The results show that the undulating surface causes flow separations and reattachments in valleys and strong crossflow along the leeward. k-ω-γ model can correctly predict both the shapes and locations of the transition onsets. The transition zone predicted by γ-Reθ model is much small, while kT-kL-ω is incapable of transition prediction for such a complex and irregular configuration. Moreover, this study reveals that the crossflow instability plays a dominant role in the transition. The crossflow Reynolds number Recf, whose distributions are approximately consistent with the transition zone, could be a feasible crossflow strength indicator for the transition onsets on the undulating surface. Once k-ω-γ model excludes effects of crossflow instability, it predicts incorrect transition results for the deformed surface of HIAD. Besides, a detailed analysis shows that both the crossflow instability mode and the first disturbance mode in valleys are the major contributors to the transition on the leeward surface. Near the leeward ray, the transition is mainly determined by the first disturbance mode.

Computational Fluid Dynamics Simulations of Supersonic Inflatable Aerodynamic Decelerator Ballistic Range Tests

The free-flight behavior of entry vehicles can greatly impact vehicle and mission design. Current methodologies rely heavily on ground and flight experiments, while computational fluid dynamics is largely used as a complement to experiments in the form of static aerodynamic databases. The dynamic stability of entry vehicles is primarily studied through ballistic range experiments and flight tests. In an effort to validate the predictive capabilities of computational fluid dynamics for free-flight aerodynamic behavior, numerical simulations of a ballistic range experiment are performed using the unstructured finite-volume Navier–Stokes solver, US3D. The ballistic range tests used for comparison in this paper were performed on a scaled model of the Supersonic Inflatable Aerodynamic Decelerator geometry. The purpose of these experiments was to provide aerodynamic coefficients of the vehicle as an a priori analysis of dynamic stability coefficients ahead of the Supersonic Flight Demonstration Test. The range data span the moderate Mach number range, 3.8–2.0, with a maximum total angle-of-attack range, 10–20 deg. These conditions are intended to span the Mach/angle-of-attack space for the majority of the flight test. Comparisons to raw vehicle attitude and postprocessed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both the raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near-body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data give further insight into dynamic/flow coupling, leading to a potential mechanism for dynamic instability.

Reconstruction of Atmosphere, Trajectory, and Aerodynamics for the Low-Density Supersonic Decelerator Project

The Supersonic Flight Dynamics Test was a full-scale flight test of aerodynamic decelerator technologies developed by the Low-Density Supersonic Decelerator technology development project. The purpose of the project was to develop and mature aerodynamic decelerator technologies for landing large mass payloads on the surface of Mars. The technologies included an attached toroidal Supersonic Inflatable Aerodynamic Decelerator, a ballute, and supersonic parachutes. Two flight tests were conducted at the Pacific Missile Range Facility. Sufficient data were acquired to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics for these two flight tests. This paper describes the instrumentation, analysis techniques, acquired data, and reconstruction results from these two flight tests. A technique for processing gimballed inertial measurement unit data is developed in which the inertial flight path of the instrument is reconstructed, which is then mapped back to the vehicle body frame as a postprocessing step.

Structural analysis of hypersonic inflatable aerodynamic decelerator pressure tub testing

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) systems have the potential to deliver the large payloads required for human-scale Mars missions. The structural response of the HIAD is critical as the inflatable members are relatively compliant compared with traditional decelerators. Structural testing and analysis were conducted on a 3.7 m HIAD subjected to uniform pressure. A beam-based finite-element modeling methodology was developed to analyze the inflatable system that incorporates strap pretension and interaction between adjacent members, a significant advance relative to previous analyses of individual inflated HIAD components. The computationally efficient modeling approach accurately captured the load-deformation response of the HIAD system.

Surface Heating and Boundary-Layer Transition on a Hypersonic Inflatable Aerodynamic Decelerator

The aerothermodynamic environment of a hypersonic inflatable aerodynamic decelerator with a flexible thermal protection system has been investigated through wind-tunnel testing. Boundary-layer transition onset locations and surface heating distributions were measured using global phosphor thermography on deflected, solid-surface models representative of a hypersonic inflatable aerodynamic decelerator aeroshell with a flexible surface. Data were obtained at Mach 6 for a wide range of surface deflection heights, angles of attack, and freestream Reynolds numbers. The surface deflections resulted in earlier boundary-layer transition onset and higher heating levels than for a smooth-wall surface, with the greatest effects occurring on the leeward side of the aeroshell. The heating augmentation and transition onset data were correlated using flowfield properties obtained from a complementary computational study. A method was developed to estimate flexible thermal protection system deflection effects using these correlations along with smooth-wall flowfield predictions. This method was validated through comparisons to the wind-tunnel data and then used to generate postflight predictions for aeroshell heating levels of the Inflatable Reentry Vehicle Experiment flight-test program. The results of this study provide a means for estimating flexible thermal protection system deflection effects for future flight tests as well as an experimental database for use in the development and validation of computational methods for simulations of hypersonic inflatable aerodynamic decelerator aerothermodynamic environments.

Inflatable structure for aerospace application: Historical perspective and future outlook

This is an overview topic of inflatable structure for aerospace application. An inflatable structure is a promising choice for a wide range of aerospace application. Among the application are inflatable wing, space antenna, solar array, inflatable aeroshell and inflatable aerodynamic decelerators. The advantages of inflatable structure are it offers the potential for compact stowing of lightweight structure. This will result in low storage volume, lightweight, efficient packaging, easy to deploy and ultimately lead to reduction in mission cost. Historically, inflatable technology has been introduced as a concept in aerospace application with a patent on flying glider that uses inflatable tubular segments connected to the fuselage. Over the years, many concepts of inflatable were introduced but only a small number of concepts have progress beyond the stage of design and even fewer manage to be completed and used extensively.

Structural testing and analysis of a braided, inflatable fabric torus structure

Inflatable structural members have military, disaster relief, aerospace and other important applications as they possess low mass, can be stored in a relatively small volume and have significant load-carrying capacity once pressurized. Of particular interest to the present research is the Hypersonic Inflatable Aerodynamic Decelerator (HIAD) structure under development by NASA. In order to make predictions about the structural response of the HIAD system, it is necessary to understand the response of individual inflatable tori composing the HIAD structure. These inflatable members present unique challenges to structural testing and modeling due to their internal inflation pressure and relative compliance. Structural testing was performed on a braided, inflatable, toroidal structural member with axial reinforcing cords. The internal inflation pressure, magnitude of enforced displacement and loading methodology were varied. In-plane and out-of-plane experimental results were compared to model predictions using a three dimensional, corotational, flexibility-based fiber-beam finite element model including geometric and material nonlinearities, as well as the effects of inflation pressure. It was found that in order to approximate the load-deformation response observed in experimentation it is necessary to carefully control the test and model boundary conditions and loading scheme.

Protection of surface assets on Mars from wind blown jettisoned spacecraft components

Jettisoned Entry, Descent and Landing System (EDLS) hardware from landing spacecraft have been observed by orbiting spacecraft, strewn over the Martian surface. Future Mars missions that land spacecraft close to prelanded assets will have to use a landing architecture that somehow minimises the possibility of impacts from these jettisoned EDLS components. Computer modelling is used here to investigate the influence of wind speed and direction on the distribution of EDLS components on the surface. Typical wind speeds encountered in the Martian Planetary Boundary Layer (PBL) were found to be of sufficient strength to blow items having a low ballistic coefficient, i.e. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) or parachutes, onto prelanded assets even when the lander itself touches down several kilometres away. Employing meteorological measurements and careful characterisation of the Martian PBL, e.g. appropriate wind speed probability density functions, may then benefit future spacecraft landings, increase safety and possibly help reduce the delta v budget for Mars landers that rely on aerodynamic decelerators.

Effect of static shape deformation on aerodynamics and aerothermodynamics of hypersonic inflatable aerodynamic decelerator

The inflatable aerodynamic decelerator (IAD), which allows heavier and larger payloads and offers flexibility in landing site selection at higher altitudes, possesses potential superiority in next generation space transport system. However, due to the flexibilities of material and structure assembly, IAD inevitably experiences surface deformation during atmospheric entry, which in turn alters the flowfield around the vehicle and leads to the variations of aerodynamics and aerothermodynamics. In the current study, the effect of the static shape deformation on the hypersonic aerodynamics and aerothermodynamics of a stacked tori Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is demonstrated and analyzed in detail by solving compressible Navier-Stokes equations with Menter's shear stress transport (SST) turbulence model. The deformed shape is obtained by structural modeling in the presence of maximum aerodynamic pressure during entry. The numerical results show that the undulating shape deformation makes significant difference to flow structure. In particular, the more curved outboard forebody surface results in local flow separations and reattachments in valleys, which consequently yields remarkable fluctuations of surface conditions with pressure rising in valleys yet dropping on crests while shear stress and heat flux falling in valleys yet rising on crests. Accordingly, compared with the initial (undeformed) shape, the corresponding differences of surface conditions get more striking outboard, with maximum augmentations of 379 pa, 2224 pa, and 19.0W/cm2, i.e., 9.8%, 305.9%, and 101.6% for the pressure, shear stress and heat flux respectively. Moreover, it is found that, with the increase of angle of attack, the aerodynamic characters and surface heating vary and the aeroheating disparities are evident between the deformed and initial shape. For the deformable HIAD model investigated in this study, the more intense surface conditions and changed flight aerodynamics are revealed, which is critical for the selection of structure material and design of flight control system.

Thermal Protection System Response Uncertainty of a Hypersonic Inflatable Aerodynamic Decelerator

The objective of this paper is to investigate the uncertainty in the bondline temperature response of a flexible thermal protection system subject to uncertain parameters in the hypersonic flowfield and thermal response modeling of a hypersonic inflatable aerodynamic decelerator configuration for ballistic Mars entry. An inflatable decelerator with a 10 m major diameter is selected for this study based on the forebody dimensions scaled from the 6 m test article that was tested in the NASA Ames Research Center’s National Full-Scale Aerodynamics Complex facility. A global nonlinear sensitivity analysis study for the bondline temperature uncertainty shows that the dimension of uncertain parameters can reduce from 22 to 8. An uncertainty analysis of the bondline temperature in the reduced dimensions indicates that the bondline temperature varies by as much as 125% above the nominal prediction and exceeds the temperature limit of 400°C. The largest uncertainty occurred at 70 s in the trajectory before separation of the inflatable decelerator for transition to a secondary descent technology. The main contributors to the bondline temperature uncertainty are the insulator and outer fabric conductivities, as well as the freestream density. The thickness and initial density of the insulator layer, closest to the gas barrier layer, are also shown to be significant contributors to the bondline temperature uncertainty, especially earlier in the trajectory.

Uncertainty Analysis of Fluid-Structure Interaction of a Deformable Hypersonic Inflatable Aerodynamic Decelerator

The objective of this paper is to present the results of a detailed uncertainty analysis for high-fidelity fluid–structure interaction modeling of a deformable hypersonic inflatable aerodynamic decelerator at peak heating conditions for lifting Mars entry with a turbulent flow assumption. Uncertainty results are presented for the structural deformation response and surface conditions (pressure, shear stress, and convective heat transfer) of the inflatable decelerator with an efficient polynomial chaos expansion approach. The uncertainty results are compared with results obtained in a previous study for ballistic Mars entry. Approximately half of the flowfield and structural modeling uncertainties show at least 90% combined contribution to the inflatable decelerator deflection and resulting surface condition uncertainties. For lifting Mars entry, global nonlinear sensitivity analysis shows that the tensile stiffness of the inflatable structure’s axial cords and radial straps and the torus torsional and tensile stiffnesses are the main contributors to the inflatable decelerator deflection uncertainty. As a result of these structural uncertainty contributions, the shape deformation contributes up to 10% of the uncertainty in the surface conditions. However, the freestream density dominates the uncertainty in the surface conditions experienced by the inflatable decelerator. In addition, the binary collision interaction is a significant contributor to aerodynamic heating and shear stress uncertainty.

Uncertainty Analysis of Radiative Heating Predictions for Titan Entry

The objective of this study was to investigate the uncertainty in shock layer radiative heating predictions on the surface of a hypersonic inflatable aerodynamic decelerator during Titan entry at peak radiative heating conditions. Computational fluid dynamics simulations of planetary entry flows and radiative heating predictions possess a significant amount of uncertainty due to the complexity of the flow physics and the difficulty in obtaining accurate experimental results of molecular-level phenomena. Sources of uncertainty considered include flowfield chemical rate models, molecular band emission, and the excitation/deexcitation rates of molecules modeled with a non-Boltzmann approach. Because of the computational cost of the numerical models, uncertainty quantification was performed with a surrogate modeling approach based on a sparse approximation of the point-collocation nonintrusive polynomial chaos expansion. Accurate uncertainty results were obtained with only 500 evaluations of the computational model. Results showed that epistemic uncertainty intervals of surface radiative heating predictions were as wide as during Titan entry, indicating the significant effect of uncertainty. A global nonlinear sensitivity analysis showed that the top uncertainty source contributing to the uncertainty in radiative heating was the flowfield chemistry modeling throughout the shock layer.

Uncertainty Analysis of Mars Entry Flows over a Hypersonic Inflatable Aerodynamic Decelerator

A detailed uncertainty analysis for high-fidelity flowfield simulations over a fixed aeroshell of hypersonic inflatable aerodynamic decelerator scale for Mars entry is presented for fully laminar and turbulent flows at peak stagnation-point heating conditions. This study implements a sparse-collocation approach based on stochastic expansions for efficient and accurate uncertainty quantification under a large number of uncertainty sources in the computational model. The convective and radiative heating and shear stress uncertainties are computed over the hypersonic inflatable aerodynamic decelerator surface and are shown to vary due to a small fraction of 65 flowfield and radiation modeling parameters considered in the uncertainty analysis. The main contributors to the convective heating uncertainty near the stagnation point are the , , and CO–O binary collision interactions, freestream density, and freestream velocity for both boundary-layer flows. In laminar flow, exothermic recombination reactions are more important at the shoulder. The main contributors to radiative heating at the nose and flank were the dissociation rate and CO heavy-particle excitation rates, whereas the freestream density showed importance toward the shoulder. The interaction and freestream velocity and density control the wall shear stress uncertainty.

Investigation of Drag-Modulated Supersonic Inflatable Aerodynamic Decelerators for Sounding Rocket Payloads

The goal of this investigation is to understand the sizing and performance of supersonic inflatable aerodynamic decelerators for Earth-based sounding rocket applications. The recovery system under examination is composed of a supersonic inflatable aerodynamic decelerator and a guided parafoil system to achieve sub-100 m miss distances. Three supersonic inflatable aerodynamic decelerator configurations (tension cone, attached isotensoid, and trailing isotensoid) are examined using the metrics of decelerator mass, aerodynamic performance, and vehicle integration. In terms of aerodynamic performance, the tension cone is the preferred choice for the sizes investigated. The attached isotensoid was shown to be the most mass efficient decelerator, whereas the trailing isotensoid was found to be the more ideal decelerator for vehicle integration. A three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo uncertainty analysis to assess the landed accuracy capability of the proposed architectures. In 95% of the cases examined, the drag-modulated inflatable aerodynamic decelerator provides arrivals within the 10 km parafoil capability region, meeting the sub-100 m landed recovery goals. In 76% of the cases examined, the drag-modulated inflatable aerodynamic decelerator arrives within 5 km of this target zone.

Conceptual Modeling of Drag-Augmented Supersonic Retropropulsion for Mars Entry, Descent, and Landing

The development of new decelerator technologies will be required as the payload mass for future Mars landing missions increases beyond the current state-of-the-art capability. This study examines the potential for supersonic retropropulsion applied on entry, descent, and landing vehicles to increase the landed payload mass. This study describes the development of a model characterizing the drag augmentation capabilities of peripheral-nozzle supersonic retropropulsion flow interactions. The model captures the dominant flow physics of pressure conservation through shock cascade structures and predicts an increase in the drag coefficient over the nominal drag coefficient of a 70 deg sphere-cone aeroshell by 14% at high Mach numbers. This study also describes drag-augmented supersonic retropropulsion operation concepts for use in Mars entry, descent, and landing. Drag-augmented supersonic retropropulsion is found to be most effective when used in the region of maximum freestream dynamic pressure. The vehicle dry mass is increased by 47% over the reference ballistic trajectory. The region of influence for aerodynamic–propulsive interactions is identified for a set of constant-thrust supersonic retropropulsion trajectories. A hybrid concept combining supersonic retropropulsion and an inflatable aerodynamic decelerator is found to be capable of providing vehicle dry masses that are 707% larger than the baseline vehicle studied.

Uncertainty Quantification of Hypersonic Reentry Flows with Sparse Sampling and Stochastic Expansions

The objective of this study was to introduce a combined sparse sampling and stochastic expansion approach for efficient and accurate uncertainty quantification. The new techniques are applied to high-fidelity, hypersonic reentry flow simulations, which may contain large numbers of aleatory and epistemic uncertainties. Stochastic expansion coefficients were obtained using the point-collocation nonintrusive polynomial chaos technique with a number of samples less than the minimum number required for a total-order expansion. This study introduced two methods of measuring the accuracy of the expansion coefficients, as well as their convergence with iteratively increasing sample size. The newly developed approaches were demonstrated on two model problems. The first was a model for stagnation point, convective heat transfer in hypersonic flow. Mixed uncertainty quantification analysis results showed that accurate expansion coefficients could be obtained with half of the minimum number of samples required for a total-order expansion. The second problem was a high-fidelity, computational fluid dynamics model for radiative heat flux prediction on a hypersonic inflatable aerodynamic decelerator during Mars entry. The model consisted of 93 uncertain parameters, coming from both flowfield and radiation modeling. Results indicated that an accurate surrogate model could be obtained with only about 15% of the number of samples required for a total-order expansion when compared with previous work.

Aeroelastic Stability of Thermal Protection System for Inflatable Aerodynamic Decelerator

A theoretical aeroelastic stability analysis has been performed on the flexible thermal protection system for an inflatable aerodynamic decelerator. Structural models consist of one or more truncated conical shells of the Donnell type, which may be elastically supported along the middle surface. The aerodynamic model is first-order piston theory. The Lagrangian of the system is formulated in terms of the generalized coordinates for all shell displacements, and the Rayleigh–Ritz method is used to derive the equations of motion. The aeroelastic stability boundaries and mode shapes are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the thermal protection system is approximated as a single conical shell, circumferentially asymmetric coalescence flutter between the second and third axial modes is observed. When many circumferential elastic supports are included, the shell flutters symmetrically in zero circumferential waves, with the first, second, and third axial modes being the most critical. In this case, the flutter boundary, flutter mechanism, and critical modes may change significantly with the addition of structural damping. Aeroelastic models that consider the thermal protection system as multiple interacting shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models, with higher axial modes being more critical. It is also found that tension applied at the shell edges, orthotropicity, and elastic support stiffness are important parameters that can dramatically affect the shell’s flutter behavior.

Methodology for Flight-Relevant Arc-Jet Testing of Flexible Thermal Protection Systems

A methodology to correlate flight aeroheating environments to the arc-jet environment is presented. For a desired hot-wall flight heating rate, the methodology provides the arc-jet bulk enthalpy for the corresponding cold-wall heating rate. A series of analyses were conducted to examine the effects of the test sample model holder geometry to the overall performance of the test sample. The analyses were compared with arc-jet test samples, and challenges and issues are presented. The transient flight environment was calculated for the Hypersonic Inflatable Aerodynamic Decelerator Earth Atmospheric Reentry Test vehicle, which is a planned demonstration vehicle using a large inflatable, flexible thermal protection system to reenter the Earth’s atmosphere from the International Space Station. A series of correlations were developed to define the relevant arc-jet test environment to properly approximate the vehicle flight environment. The computed arc-jet environments were compared with the measured arc-jet values to define the uncertainty of the correlated environment. The results show that, for a given flight surface heat flux and a fully catalytic thermal protection system, the flight-relevant arc-jet heat flux increases with the arc-jet bulk enthalpy, while for a noncatalytic thermal protection system, the arc-jet heat flux decreases with the bulk enthalpy.