Supersonic Decelerator Research Articles

Advanced Supersonic Parachute Inflation Research Experiment Preflight Trajectory Modeling and Postflight Reconstruction

The Advanced Supersonic Parachute Inflation Research Experiment (ASPIRE) was a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars robotic applications. Although the 2012 Mars Science Laboratory (MSL) had a successful deployment of a supersonic parachute, postflight analysis and the results of the Low-Density Supersonic Decelerator project called into question some of the assumptions underlying MSL’s parachute qualification program. The ASPIRE project was a key part of the risk-reduction program undertaken by Mars 2020, the follow-on to MSL. The first sounding rocket (SR01) flight of ASPIRE occurred near Wallops Island, Virginia on 4 October 2017, and demonstrated the successful deployment and inflation of an MSL build-to-print parachute in flight conditions similar to the 2012 MSL mission. ASPIRE SR02 and SR03 were successful follow-on flights on 31 March 2018 and 7 September 2018 that demonstrated the new, strengthened supersonic parachute designed for the Mars 2020 project. The SR02 and SR03 parachutes were targeted to 100 and 140% of the expected flight limit load for Mars 2020 to confirm new margins expected from the strengthened parachute. Prior to all flights, a multibody flight dynamics simulation was developed to predict the parachute dynamics and was used, in conjunction with other tools, to target Mars-relevant flight conditions. After each flight, the onboard data were used to reconstruct the flight trajectory and to validate the preflight dynamics simulation. Postflight analysis showed that all three tests achieved their targeted conditions and preflight modeling bounded the key performance metrics for the parachute. This paper describes the flight mechanics simulation, postflight reconstruction, and the reconciliation process used to validate the flight models.

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.

Fluid–Structure Interaction Simulations of a Tension-Cone Inflatable Aerodynamic Decelerator

A series of fluid–structure interaction simulations of an aerodynamic tension-cone supersonic decelerator prototype intended for large mass payload deployment in planetary explorations are discussed. The fluid–structure interaction computations combine large deformation analysis of thin shells with large-eddy simulation of compressible turbulent flows using a loosely coupled approach to enable quantification of the dynamics of the vehicle. The simulation results are compared with experiments carried out at the NASA Glenn Research Center. Reasonably good agreement between the simulations and the experiment is observed throughout a deflation cycle. The simulations help to illuminate the details of the dynamic progressive buckling of the tension-cone decelerator that ultimately results in the collapse of the structure as the inflation pressure is decreased. Furthermore, the tension-cone decelerator exhibits a transient oscillatory behavior under impulsive loading that ultimately dies out. The frequency of these oscillations was determined to be related to the acoustic time scale in the compressed subsonic region between the bow shock and the structure. As shown, when the natural frequency of the structure and the frequency of the compressed subsonic region approximately match, the decelerator exhibits relatively large nonaxisymetric oscillations. The observed response appears to be a fluid–structure interaction resonance resulting from an acoustic chamber (pistonlike) mode exciting the structure.

A concept for the entry, descent, and landing of high-mass payloads at Mars

The architecture concepts and aggressive science objectives for the next phases of Mars exploration will require landed masses an order of magnitude or greater than any Mars mission previously planned or flown. Additional studies have shown the requirements for missions more ambitious than the 2009 Mars Science Laboratory (∼900 kg payload mass) to extend beyond the capabilities of Viking-heritage entry, descent, and landing (EDL) technologies, namely blunt-body aeroshells, supersonic disk-gap-band parachutes, and existing TPS materials. This study details a concept for Mars entry, descent, and landing capable of delivering a 20 t payload within 1 km of a target landing site at 0 km MOLA. The concept presented here explores potentially enabling EDL technologies for the continued robotic and eventual human exploration of Mars, moving beyond the Viking-heritage systems relied upon for the past 30 years of Mars exploration. These technologies address the challenges of hypersonic guidance, supersonic deceleration, precision landing, and surface hazard avoidance. Without support for the development of these enabling technologies in the near term, the timeline for the successful advanced exploration of Mars will likely extend indefinitely.

Problem of Deceleration of a Supersonic Conducting Channel Flow by a Magnetic Field

Problems of the deceleration of a supersonic conducting flow by a magnetic field are investigated. A conducting gas flow in a circular tube is considered in the presence of an axisymmetric magnetic field induced by a unit current loop or solenoid of finite length. The analysis is carried out on the basis of both the Euler equations (inviscid gas) and the complete system of Navier-Stokes equations for laminar viscous gas flow and turbulent flow using a one-parameter turbulence model. The numerical simulation is based on an implicit relaxation finite-difference scheme which is a modification of the Godunov method. The total pressure losses are determined for various values of the magnetohydrodynamic (MHD) interaction, the initial Mach number, and different magnetic field geometries and it is shown that the irreversible losses are significant in MHD supersonic flow deceleration.

Supersonic flow deceleration in plane and axisymmetric channels

Supersonic perfect gas flow in plane and axisymmetric channels with the same duct contour is studied on the basis of a numerical solution of the two-dimensional Navier-Stokes and Euler equations. The calculations were carried out at an inlet Mach number M∞=4 for various Reynolds numbers and “bell-mouth“ half-angles. The effect of these parameters, as well as that of the flow three-dimensionality, on the flow pattern is demonstrated. In particular, the existence of viscous flow regimes providing the most effective supersonic flow deceleration and a higher degree of total pressure recovery as compared with the inviscid flow is established.

Investigation on onset of shock-induced separation

A great number of experimental data indicating shock wave/boundary layer interactions in internal or external supersonic flows were reviewed to make clear the mechanism of the interaction and to decide the onset of shock-induced separation. The interesting conclusions were obtained for the considerably wide range of flow geometries that the onset of separation is independent of the flow geometries and the boundary layer Reynolds number. It is found that the pressure rise necessary to separate the boundary layer in supersonic external flows could be applied to such internal flows as overexpanded nozzles or diffusers. This is due to the fact that the separation phenomenon caused by shock wave/boundary layer interactions is processed through a supersonic deceleration. The shock-induced separation in almost all of interacting flow fields is governed by the concept of free interaction, and the onset of shock-induced separation is only a function of the Mach number just upstream of shock wave. However, physical scales of the produced separation are not independent of the downstream flow fields.

Inviscid Supersonic Far Wake Flow past Pointed Bodies

Theme T flowfield analysis of supersonic decelerators towed in the wake of fast-moving re-entry vehicles is one of significant import to predict their performance characteristics. Since many supersonic decelerators, such as inflatable ballutes and guide surface parachutes, have conical front ends to guide the flow and stabilize the shock ahead of them, simple geometrical configurations such as cones and wedges are considered in the present analysis. The method of integral relations hitherto applied to blunt bodies in uniform flows is extended to nonuniform far wake flow past pointed bodies. The analysis encompasses both attached and detached shock situations at the apex and includes the transverse variations of wake flow properties and their derivatives ahead of the shock.

A Contribution to the Theoretical and Experimental Examination of the Flow Through Plane Supersonic Deceleration Cascades and Supersonic Compressor Rotors

The results of the experimental investigations of plane supersonic deceleration cascades (tandem cascades) conducted in recent years at the Institut fu¨r Strahlantriebe und Turboarbeitsmaschinen der Technischen Hochschule Aachen (Institute for Jet Propulsion and Turbine Machines of the Aachen Technical University) are summarized. The flow through a supersonic compressor rotor is determined by means of a suitably enlarged method of characteristics. A division of the blade channel into two sets of stream surfaces permits an iterative calculation of the three-dimensional flow. To determine the flow values in the subsonic range, 5-hole-probes featuring a semi-spherical head are used in front of the rotor. The supersonic flow is determined by means of 5-hole-probes of the conical type arranged at varying distances behind the rotor. To determine the efficiencies, all probes were equipped with two NTC-elements each, for temperature measurement. These measurements were carried out at different rotational speeds and throttle positions and supplied rotor characteristics as well as information on the course of the flow behind the rotor. With unthrottled flow, a total pressure ratio of 3.9 resulted, with a total isentropic efficiency of 89 percent. The measurement of wall pressures on hub and tip furnishes additional information on the course of flow. The test values measured show a satisfactory agreement with the calculated design values.