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Deceleration, aerodynamic heating and weight factors in feasibility study of hot air balloon recovery system for large boosters
An inflatable aerodynamic deceleration method and system is provided for use with an atmospheric entry payload. The inflatable aerodynamic decelerator includes an inflatable envelope and an inflatant, wherein the inflatant is configured to fill the inflatable envelope to an inflated state such that the inflatable envelope surrounds the atmospheric entry payload, causing aerodynamic forces to decelerate the atmospheric entry payload.
Two inflatable aerodynamic decelerator designs were tested in the Transonic Dynamics Tunnel at the NASA Langley Research Center: a tension cone and an isotensoid. The tension cone consists of a flexible tension shell attached to a torus and the isotensoid employs a ram-air inflated envelope. Tests were conducted at Mach numbers from 0.3 to 1.08 and Reynolds numbers from 0.59 to 2.46 million. The main objective of these tests was to obtain static aerodynamic coefficients at subsonic and transonic speeds to supplement supersonic aerodynamic data for these same two designs. The axial force coefficients of both designs increased smoothly from subsonic through transonic Mach numbers. Dynamic data show significant oscillation of the tension cone and minimal oscillation of the isotensoid. The transonic and subsonic data will be used to assemble an inflatable decelerator aerodynamic database for use in computational analyses and system studies.
In this paper, aerodynamic decelerators are defined as textile devices intended to be deployed at Mach numbers below five. Such aerodynamic decelerators include parachutes and inflatable aerodynamic decelerators (often known as ballutes). Aerodynamic decelerators play a key role in the Entry, Descent, and Landing (EDL) of planetary exploration vehicles. Among the functions performed by aerodynamic decelerators for such vehicles are deceleration (often from supersonic to subsonic speeds), minimization of descent rate, providing specific descent rates (so that scientific measurements can be obtained), providing stability (drogue function - either to prevent aeroshell tumbling or to meet instrumentation requirements), effecting further aerodynamic decelerator system deployment (pilot function), providing differences in ballistic coefficients of components to enable separation events, and providing height and timeline to allow for completion of the EDL sequence. Challenging aspects in the development of aerodynamic decelerators for planetary exploration missions include: deployment in the unusual combination of high Mach numbers and low dynamic pressures, deployment in the wake behind a blunt-body entry vehicle, stringent mass and volume constraints, and the requirement for high drag and stability. Furthermore, these aerodynamic decelerators must be qualified for flight without access to the exotic operating environment where they are expected to operate. This paper is an introduction to the development and application of aerodynamic decelerators for robotic planetary exploration missions (including Earth sample return missions) from the earliest work in the 1960s to new ideas and technologies with possible application to future missions. An extensive list of references is provided for additional study.
The supersonic aerodynamic and aeroelastic characteristics of a tension cone inflatable aerodynamic decelerator were investigated by wind tunnel testing. Two sets of tests were conducted: one using rigid models and another using textile models. Tests using rigid models were conducted over a Mach number range from 1.65 to 4.5 at angles of attack from -12 to 20 degrees. The axial, normal, and pitching moment coefficients were found to be insensitive to Mach number over the tested range. The axial force coefficient was nearly constant (C(sub A) = 1.45 +/- 0.05) with respect to angle of attack. Both the normal and pitching moment coefficients were nearly linear with respect to angle of attack. The pitching moment coefficient showed the model to be statically stable about the reference point. Schlieren images and video showed a detached bow shock with no evidence of large regions of separated flow and/or embedded shocks at all Mach numbers investigated. Qualitatively similar static aerodynamic coefficient and flow visualization results were obtained using textile models at a Mach number of 2.5. Using inflatable textile models the torus pressure required to maintain the model in the fully-inflated configuration was determined. This pressure was found to be sensitive to details in the structural configuration of the inflatable models. Additional tests included surface pressure measurements on rigid models and deployment and inflation tests with inflatable models.
Supersonic inflatable aerodynamic decelerators (IADs) are drag devices intended to be deployed at high Mach numbers. In the application considered here they assist in the descent and landing of spacecraft on Mars. Although promising, present IAD technology is not yet sufficiently mature for use in the near future. This paper describes a technology maturation plan for tension cone IADs using subscale test articles to reduce development costs. As envisioned, the proposed test plan includes three phases: wind tunnel tests (subsonic), unpowered high-altitude flight tests (transonic), and powered high-altitude tests (supersonic). This test plan is based on a building block approach in which successful completion of each phase adds to the understanding of the behavior of IADs and reduces the risk of the subsequent, more expensive phases. By properly scaling the IADs, test articles of the same size and nearly the same construction can be used for all three phases. The final phase is a dynamically scaled flight test with IAD deployment at the same Mach number as the full-scale vehicle on Mars. Two full-scale example cases are presented: one for a single-stage system (15 m dia. IAD to subsonic retropropulsion), and another for a two-stage system (10.5 m dia. IAD to subsonic parachute). Using scale factors of 0.333 and 0.476 yield subscale test IADs of 5 m dia. The dynamically scaled powered flight test starts at Mach 4 and an altitude of 33.5 km. Existing balloons and rocket motors are shown to be adequate to meet the required test conditions.
Aerodynamic deceleration systems for space missions, considering deployment of subsonic parachute and inflatable configuration following atmospheric entry
Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology is currently being considered for multiple atmospheric entry applications as the limitations of traditional entry vehicles have been reached. The Inflatable Re-entry Vehicle Experiment (IRVE) has successfully demonstrated this technology as a viable candidate with a 3.0 m diameter vehicle sub-orbital flight. To further this technology, large scale HIADs (6.0 8.5 m) must be developed and tested. To characterize the performance of large scale HIAD technology new instrumentation concepts must be developed to accommodate the flexible nature inflatable aeroshell. Many of the concepts that are under consideration for the HIAD FY12 subsonic wind tunnel test series are discussed below.
The Hypersonic Inflatable Aerodynamic Decelerators (HIAD) project has invested in development of multiple thermal protection system (TPS) candidates to be used in inflatable, high downmass, technology flight projects. Flexible TPS is one element of the HIAD project which is tasked with the research and development of the technology ranging from direct ground tests, modelling and simulation, characterization of TPS systems, manufacturing and handling, and standards and policy definition. The intent of flexible TPS is to enable large deployable aeroshell technologies, which increase the drag performance while significantly reducing the ballistic coefficient of high-mass entry vehicles. A HIAD requires a flexible TPS capable of surviving aerothermal loads, and durable enough to survive the rigors of construction, handling, high density packing, long duration exposure to extrinsic, in-situ environments, and deployment. This paper provides a comprehensive overview of key work being performed within the Flexible TPS element of the HIAD project. Included in this paper is an overview of, and results from, each Flexible TPS research and development activity, which includes ground testing, physics-based thermal modelling, age testing, margins policy, catalysis, materials characterization, and recent developments with new TPS materials.
Aerodynamic deceleration systems for space missions, considering deployment of subsonic parachute and inflatable configuration following atmospheric entry
Structural merit function for aerodynamic decelerators
Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology readiness levels have been incrementally matured by NASA over the last thirteen years, with most recent support from NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). Recently STMD GCDP has authorized funding and support through fiscal year 2015 (FY15) for continued HIAD ground developments which support a Mars Entry, Descent, and Landing (EDL) study. The Mars study will assess the viability of various EDL architectures to enable a Mars human architecture pathfinder mission planned for mid-2020. At its conclusion in November 2014, NASA's first HIAD ground development effort had demonstrated success with fabricating a 50 W/cm2 modular thermal protection system, a 400 C capable inflatable structure, a 10-meter scale aeroshell manufacturing capability, together with calibrated thermal and structural models. Despite the unquestionable success of the first HIAD ground development effort, it was recognized that additional investment was needed in order to realize the full potential of the HIAD technology capability to enable future flight opportunities. The second HIAD ground development effort will focus on extending performance capability in key technology areas that include thermal protection system, lifting-body structures, inflation systems, flight control, stage transitions, and 15-meter aeroshell scalability. This paper presents an overview of the accomplishments under the baseline HIAD development effort and current plans for a follow-on development effort focused on extending those critical technologies needed to enable a Mars Pathfinder mission.
Typical entry vehicle aeroshells are limited in size by the launch vehicle shroud. Inflatable aerodynamic decelerators allow larger aeroshell diameters for entry vehicles because they are not constrained to the launch vehicle shroud diameter. During launch, the hypersonic inflatable aerodynamic decelerator (HIAD) is packed in a stowed configuration. Prior to atmospheric entry, the HIAD is deployed to produce a drag device many times larger than the launch shroud diameter. The large surface area of the inflatable aeroshell provides deceleration of high-mass entry vehicles at relatively low ballistic coefficients. Even for these low ballistic coefficients there is still appreciable heating, requiring the HIAD to employ a thermal protection system (TPS). This TPS must be capable of surviving the heat pulse, and the rigors of fabrication handling, high density packing, deployment, and aerodynamic loading. This paper provides a comprehensive overview of flexible TPS tests and results, conducted over the last three years. This paper also includes an overview of each test facility, the general approach for testing flexible TPS, the thermal analysis methodology and results, and a comparison with 8-foot High Temperature Tunnel, Laser-Hardened Materials Evaluation Laboratory, and Panel Test Facility test data. Results are presented for a baseline TPS layup that can withstand a 20 W/cm2 heat flux, silicon carbide (SiC) based TPS layup, and polyimide insulator TPS layup. Recent work has focused on developing material layups expected to survive heat flux loads up to 50 W/cm2 (which is adequate for many potential applications), future work will consider concepts capable of withstanding more than 100 W/cm2 incident radiant heat flux. This paper provides an overview of the experimental setup, material layup configurations, facility conditions, and planned future flexible TPS activities.
This paper describes a technique for estimating mass for inflatable aerodynamic decelerators. The technique uses dimensional analysis to identify a set of dimensionless parameters for inflation pressure, mass of inflation gas, and mass of flexible material. The dimensionless parameters enable scaling of an inflatable concept with geometry parameters (e.g., diameter), environmental conditions (e.g., dynamic pressure), inflation gas properties (e.g., molecular mass), and mass growth allowance. This technique is applicable for attached (e.g., tension cone, hypercone, and stacked toroid) and trailing inflatable aerodynamic decelerators. The technique uses simple engineering approximations that were developed by NASA in the 1960s and 1970s, as well as some recent important developments. The NASA Mars Entry and Descent Landing System Analysis (EDL-SA) project used this technique to estimate the masses of the inflatable concepts that were used in the analysis. The EDL-SA results compared well with two independent sets of high-fidelity finite element analyses.
Verification analyses were conducted on membrane structures pertaining to a tension cone inflatable aerodynamic decelerator using the analysis code LS-DYNA. The responses of three structures - a cylinder, torus, and tension shell - were compared against linear theory for various loading cases. Stress distribution, buckling behavior, and wrinkling behavior were investigated. In general, agreement between theory and LS-DYNA was very good for all cases investigated. These verification cases exposed the important effects of using a linear elastic liner in membrane structures under compression. Finally, a tension cone wind tunnel test article is modeled in LS-DYNA for which preliminary results are presented. Unlike data from supersonic wind tunnel testing, the segmented tension shell and torus experienced oscillatory behavior when subjected to a steady aerodynamic pressure distribution. This work is presented as a work in progress towards development of a fluid-structures interaction mechanism to investigate aeroelastic behavior of inflatable aerodynamic decelerators.
Nonuniform free stream supersonic flow past aerodynamic decelerators, calculating inviscid flow fields by method of characteristics
Hypersonic Inflatable Aerodynamic Decelerators technology has been developed by NASA to enable challenging entry, descent, and landing missions at various planetary destinations. In the last two decades, HIAD technology has been developed through a series of ground tests and flight tests, culminating in the 2022 Low-Earth Orbit Flight Test of an Inflatable Decelerator flight demonstration of a 6 m diameter vehicle from orbital velocities. Past studies have considered application of HIAD for human-scale missions, especially return from the Moon or human-class missions to Mars. However, recent commercial interest in the HIAD technology following LOFTID’s success shows HIAD applications are possible for cislunar return, Earth aerocapture, return from low Earth orbit, and even launch vehicle component recovery. This paper describes the flight performance range for these new, Earth-based applications.
Hypersonic Inflatable Aerodynamic Decelerators technology has been developed by NASA to enable challenging entry, descent, and landing missions at various planetary destinations. In the last two decades, HIAD technology has been developed through a series of ground tests and flight tests, culminating in the 2022 Low-Earth Orbit Flight Test of an Inflatable Decelerator flight demonstration of a 6 m diameter vehicle from orbital velocities. Past studies have considered application of HIAD for human-scale missions, especially return from the Moon or human-class missions to Mars. However, recent commercial interest in the HIAD technology following LOFTID’s success shows HIAD applications are possible for cislunar return, Earth aerocapture, return from low Earth orbit, and even launch vehicle component recovery. This paper describes the flight performance range for these new, Earth-based applications.