Results for “inflation loads” — Page 11

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.

Supersonic Retropropulsion↗

NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testing on a Blunt Mars Entry Vehicle Concept

Supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle was simulated using the Overflow Computational Fluid Dynamics (CFD) solver. Two nozzle configurations were tested (1E and 1F) as a subset of the seven total configurations in the Descent System Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and unique grid systems optimized for any given testing condition. Solution independence from grid resolution was determined by successively increasing grid refinement until asymptotic convergence of the mean aerodynamic loads was observed. Overflow Reynolds-averaged Navier-Stokes (RANS) solutions produced realistic SRP flow phenomena, including Mach disk normal shocks contained in high-thrust plumes and bow shock triple-points. Solutions for the 1E nozzle configuration were steady. Conversely, a subset of the 1F test conditions were unsteady and exhibited periodic, non-sinusoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle. Initial investigations into the effect of turbulence modeling fidelity demonstrated significant differences between the aerodynamic loads simulated with RANS and Detached Eddy Simulation (DES). Multiple flow mechanisms were identified as root causes of these differences, including bow shock shape augmentation and reduced entrainment due to reduced turbulence in the DES simulations. Initial comparisons of pitching moment control authority and aerodynamic drag performance between the tested nozzle configurations demonstrated that the 1E configuration may be more advantageous for the Mars Entry, Descent, and Landing (EDL) task.

Supersonic Retropropulsion↗

NASA NTRS1972

Static structural tests of a 1.5-meter diameter fabric attached inflatable decelerator

Meridional tape forces, permeability, and change in model contours were measured on the aft half of a 1.5-m-diameter attached inflatable decelerator (AID). Inflation pressures up to 103 kN per square meter and external pressures of 0.14 to 102 kN per square meter were used. The results indicated that the model stresses were near the desired isotensoid condition. Future AID designs should consider both stiffness and strength of the meridional tapes to obtain the optimum division of load between tapes and fabric. Permeability of pressurized fabric structures is a few orders of magnitude higher than that obtained in the standard low-pressure test on material specimens.

Willis, C. M.↗

NASA NTRS2013

Aerodynamic Reconstruction Applied to Parachute Test Vehicle Flight Data Analysis

The aerodynamics, both static and dynamic, of a test vehicle are critical to determining the performance of the parachute cluster in a drop test and for conducting a successful test. The Capsule Parachute Assembly System (CPAS) project is conducting tests of NASA's Orion Multi-Purpose Crew Vehicle (MPCV) parachutes at the Army Yuma Proving Ground utilizing the Parachute Test Vehicle (PTV). The PTV shape is based on the MPCV, but the height has been reduced in order to fit within the C-17 aircraft for extraction. Therefore, the aerodynamics of the PTV are similar, but not the same as, the MPCV. A small series of wind tunnel tests and computational fluid dynamics cases were run to modify the MPCV aerodynamic database for the PTV, but aerodynamic reconstruction of the flights has proven an effective source for further improvements to the database. The acceleration and rotational rates measured during free flight, before parachute inflation but during deployment, were used to con rm vehicle static aerodynamics. A multibody simulation is utilized to reconstruct the parachute portions of the flight. Aerodynamic or parachute parameters are adjusted in the simulation until the prediction reasonably matches the flight trajectory. Knowledge of the static aerodynamics is critical in the CPAS project because the parachute riser load measurements are scaled based on forebody drag. PTV dynamic damping is critical because the vehicle has no reaction control system to maintain attitude - the vehicle dynamics must be understood and modeled correctly before flight. It will be shown here that aerodynamic reconstruction has successfully contributed to the CPAS project.

Cassady, Leonard D.↗

NASA NTRS1974

Deployment and performance characteristics of 1.5-meter supersonic attached inflatable decelerators

Attached-inflatable-decelerator (AID) canopies fabricated from lightweight Nomex cloth and tapes were deployed in a supersonic stream from the base of a 140 deg conical aeroshell. Characteristics of the deceleration system were obtained over a wide range of Mach number, dynamic pressure, and pitch angle. All models deployed rapidly by ram air and experienced only mild deployment shock loads. Steady-state drag coefficients as high as 1.3 were obtained in the supersonic stream and were relatively insensitive to Mach number, dynamic pressure, and pitch angle. All models were free of fluttering motion. Results also showed that the AID is aerodynamically more efficient without a burble fence in a supersonic stream. Though measured meridian-tape loads were higher than those predicted by theory, the ram-air deployment rates and steady-state drag coefficients were in good agreement with theory. These results indicate that the AID is a stable, efficient decelerator in a supersonic stream and its performance is readily predictable.

Bohon, H. L.↗

NASA NTRS2011

Advanced High-Temperature Flexible TPS for Inflatable Aerodynamic Decelerators

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.

DelCorso, Joseph A.↗

NASA NTRS2016

Optimization of a Hot Structure Aeroshell and Nose Cap for Mars Atmospheric Entry

The National Aeronautics and Space Administration (NASA) is preparing to send humans beyond Low Earth Orbit and eventually to the surface of Mars. As part of the Evolvable Mars Campaign, different vehicle configurations are being designed and considered for delivering large payloads to the surface of Mars. Weight and packing volume are driving factors in the vehicle design, and the thermal protection system (TPS) for planetary entry is a technology area which can offer potential weight and volume savings. The feasibility and potential benefits of a ceramic matrix composite hot structure concept for different vehicle configurations are explored in this paper, including the nose cap for a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and an aeroshell for a mid lift-to-drag (Mid L/D) concept. The TPS of a planetary entry vehicle is a critical component required to survive the severe aerodynamic heating environment during atmospheric en- try. The current state-of-the-art is an ablative material to protect the vehicle from the heat load. The ablator is bonded to an underlying structure, which carries the mechanical loads associated with entry. The alternative hot structure design utilizes an advanced carbon-carbon material system on the outer surface of the vehicle, which is exposed to the severe heating and acts as a load carrying structure. The preliminary design using the hot structure concept and the ablative concept is determined for the spherical nose cap of the HIAD entry vehicle and the aeroshell of the Mid L/D entry vehicle. The results of the study indicate that the use of hot structures for both vehicle concepts leads to a feasible design with potential weight and volume savings benefits over current state-of-the-art TPS technology that could enable future missions.

Langston, Sarah L.↗

NASA NTRSDate not supplied

Overview and Performance of the LOFTID Instrumentation Suite

On November 10, 2022, NASA launched the Low-earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle as a secondary payload mounted inside the launch vehicle adaptor on an Atlas V out of the Vandenburg Space Force Base (VSFB). The primary payload, Joint Polar Satellite System-2 (JPSS-2), was delivered successfully to a sun synchronous trajectory shortly after launch, at which point the Centaur upper stage performed a burn to de-orbit the system. Once on the desired trajectory to enter the atmosphere the payload adaptor canister was ejected by the payload adapter separation system to expose the packed LOFTID vehicle, then the LOFTID aeroshell, a Hypersonic Inflatable Aerodynamic Decelerator (HIAD), was deployed and inflated as planned. The Centaur pointed the LOFTID vehicle to the desired attitude to enter the atmosphere, spun the assembly to roughly three rpm, and separated the reentry vehicle. The LOFTID vehicle entered the atmosphere over Alaska at >8km/sec and decelerated as designed demonstrating stable flight from hypersonic entry through subsonic parachute deployment. On-board visible light cameras captured the reactions of the heatshield through all phases of flight, and co-located infrared light cameras captured the temperature distribution of the aft side of the heat shield anchored to a distribution of thermocouples on the inflatable structure (IS) in the field of view. Thermocouples were also embedded in the forward side of the aeroshell both in the Flexible Thermal Protection System (FTPS) as well as the IS. Heat Flux Gages and Pressure Transducers measured the heating rate and surface pressure experienced by the rigid nose of the reentry vehicle. Loadcells measured the interface loads between the IS and the rigid centerbody. This paper will discuss many of the key instruments flown on the successful LOFTID technology demonstration mission and will provide some high level results, while pointing to more detailed papers on the post-flight analyses.

Gregory T Swanson↗

NASA NTRSDate not supplied

Overview and Performance of the LOFTID Instrumentation Suite

On November 10, 2022, NASA launched the Low-earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle as a secondary payload mounted inside the launch vehicle adaptor on an Atlas V out of the Vandenburg Space Force Base (VSFB). The primary payload, Joint Polar Satellite System-2 (JPSS-2), was delivered successfully to a sun synchronous trajectory shortly after launch, at which point the Centaur upper stage performed a burn to de-orbit the system. Once on the desired trajectory to enter the atmosphere the payload adaptor canister was ejected by the payload adapter separation system to expose the packed LOFTID vehicle, then the LOFTID aeroshell, a Hypersonic Inflatable Aerodynamic Decelerator (HIAD), was deployed and inflated as planned. The Centaur pointed the LOFTID vehicle to the desired attitude to enter the atmosphere, spun the assembly to roughly three rpm, and separated the reentry vehicle. The LOFTID vehicle entered the atmosphere over Alaska at >8km/sec and decelerated as designed demonstrating stable flight from hypersonic entry through subsonic parachute deployment. On-board visible light cameras captured the reactions of the heatshield through all phases of flight, and co-located infrared light cameras captured the temperature distribution of the aft side of the heat shield anchored to a distribution of thermocouples on the inflatable structure (IS) in the field of view. Thermocouples were also embedded in the forward side of the aeroshell both in the Flexible Thermal Protection System (FTPS) as well as the IS. Heat Flux Gages and Pressure Transducers measured the heating rate and surface pressure experienced by the rigid nose of the reentry vehicle. Loadcells measured the interface loads between the IS and the rigid centerbody. This paper will discuss many of the key instruments flown on the successful LOFTID technology demonstration mission and will provide some high level results, while pointing to more detailed papers on the post-flight analyses.

spacecraft design↗

NASA NTRS1999

Special Considerations in Selection of Fabric Film Laminates for Use in Inflatable Structures

Inflatable structures are gaining wide support in planetary scientific missions as well as commercial applications. For such applications a new class of fabric/film laminates is being considered for use as a structural gas envelope. The emerging composite materials are a result of recent advances in the manufacturing of lightweight, high strength fibers, fabrics and scrims. The lamination of these load-carrying members with the proper gas barriers results in a wide range of materials suitable for various loading and environmental conditions. Polyester-based woven fabrics laminated to thin homogenous film of polyester are an example of this class. This fabric/film laminate is being considered for the development of a material suitable for building large gas envelopes for use in the NASA ultra long duration balloon program (ULDB). Compared to commercial homogenous films, the material provides relatively high strength to weight ratio as well as better resistance to crack and tear propagation, The mechanical, creep and viscoelastic properties of these fabric film laminates have been studied to form a material model. Preliminary analysis indicates that the material is highly viscoelastic. The mechanical properties of this class of materials will be discussed in some details.

Said, Magdi A.↗

NASA NTRS2008

Excavation on the Moon: Regolith Collection for Oxygen Production and Outpost Site Preparation

The development of a robust regolith moving system for lunar and planetary processing and construction is critical to the NASA mission to the Moon and Mars. Oxygen production may require up to 200 metric tons of regolith collection per year; outpost site development may require several times this amount. This paper describes progress in the small vehicle implement development and small excavation system development. Cratos was developed as a platform for the ISRU project to evaluate the performance characteristics of a low center of gravity, small (0.75m x 0.75m x 0.3m), low-power, tracked vehicle performing excavation, load, haul, and dump operations required for lunar ISRU. It was tested on loose sand in a facility capable of producing level and inclined surfaces, and demonstrated the capability to pick up, carry, and dump sand, allowing it to accomplish the delivery of material to a site. Cratos has demonstrated the capability to pick up and deliver simulant to a bury an inflatable habitat, to supply an oxygen production plant, and to build a ramp.

Caruso, John J.↗

NASA NTRS1975

A new concept for airship mooring and ground handling

Calculations were made to determine the feasibility of applying the negative air cushion (NAC) principle to the mooring of airships. Pressures required for the inflation of the flexible trunks are not excessive and the maintenance of sufficient hold down force is possible in winds up to 50 knots. Fabric strength requirements for a typical NAC sized for a 10-million cubic foot airship were found to be approximately 200 lbs./in. Corresponding power requirements range between 66-HP and 5600-HP. No consideration was given to the internal airship loads caused by the use of a NAC and further analysis in much greater detail is required before this method could be applied to an actual design, however, the basic concept appears to be sound and no problem areas of a fundamental nature are apparent.

Vaughan, J. C.↗

NASA NTRS1999

Mechanical Behavior of Fabric-Film Laminates

Inflatable structures are gaining wide support in planetary scientific missions as well as commercial applications. For such applications a new class of materials made of laminating thin homogenous films to lightweight fabrics are being considered us structura1 gas envelops. The emerging composite materials are a result of recent advances in the manufacturing cf 1ightweight, high strength fibers, fabrics and scrims. The lamination of these load-carrying members with the proper gas barrier film results in wide range of materials suitable for various loading and environmental conditions. Polyester - based woven fabrics laminated to thin homogeneus film of polyester (Maylar) is an example of this class. This fabric/ film laminate is being considered for the development a material suitable for building large gas envelopes for use in the NASA Ultra Long Duration Balloon Program (ULDB). Compared to commercial homogeneus films, the material provides relatively high strength to weight ratio as well as better resistance to crack and tear propagation. The purpose of this papers is to introduce the mechanical behavior of this class of multi-layers composite and to highlight some of the concerns observed during the characterization of these laminate composites.

Said, Magdi S.↗

NASA NTRS2020

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

The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) was a series of sounding rocket flights aimed at understanding the dynamics of supersonic parachutes that are used for Mars robotic applications. The 2012 Mars Science Laboratory (MSL) had a successful deployment of a supersonic parachute, but based on post-flight analysis of parachute margins, the ASPIRE project was created as a risk-reduction program to improve quantification of these margins and qualify a supersonic parachute for Mars 2020, the follow-on mission to MSL. The first sounding rocket (SR01) flight of ASPIRE occurred near Wallops Island, Virginia on Oct. 4, 2017 and demonstrated the successful deployment and inflation of a MSL build-to-print parachute in flight conditions similar to the 2012 MSL mission. The ASPIRE SR02 and SR03 were successful follow-on flights on Mar. 31, 2018 and Sep. 7, 2018 that demonstrated the new, strengthened supersonic parachute designed for the Mars 2020 project. The SR02 and SR03 parachuteswere 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 multi-body 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 on-board data were used to reconstruct the flight trajectory and to validate the pre-flight dynamics simulation. Post-flight analysis showed that all three tests achieved their targeted conditions and pre-flight modeling bounded the key performance metrics for the parachute. This paper describes the flight mechanics simulation, post-flight reconstruction, and the reconciliation process used to validate the flight models.

Soumyo Dutta↗

NASA NTRS2002

Reliability Testing of NASA Piezocomposite Actuators

NASA Langley Research Center has developed a low-cost piezocomposite actuator which has application for controlling vibrations in large inflatable smart space structures, space telescopes, and high performance aircraft. Tests show the NASA piezocomposite device is capable of producing large, directional, in-plane strains on the order of 2000 parts-per-million peak-to-peak, with no reduction in free-strain performance to 100 million electrical cycles. This paper describes methods, measurements, and preliminary results from our reliability evaluation of the device under externally applied mechanical loads and at various operational temperatures. Tests performed to date show no net reductions in actuation amplitude while the device was moderately loaded through 10 million electrical cycles. Tests were performed at both room temperature and at the maximum operational temperature of the epoxy resin system used in manufacture of the device. Initial indications are that actuator reliability is excellent, with no actuator failures or large net reduction in actuator performance.

Wilkie, W.↗

NASA NTRS2011

Femoral Blood Flow and Cardiac Output During Blood Flow Restricted Leg Press Exercise

Low load blood flow restricted resistance exercise (LBFR) causes muscle hypertrophy that may be stimulated by the local ischemic environment created by the cuff pressure. However, local blood flow (BF) during such exercise is not well understood. PURPOSE: To characterize femoral artery BF and cardiac output (CO) during leg press exercise (LP) performed at a high load (HL) and low load (LL) with different levels of cuff pressure. METHODS: Eleven subjects (men/women 4/7, age 31.4+/-12.8 y, weight 68.9+/-13.2 kg, mean+/-SD) performed 3 sets of supine left LP to fatigue with 90 s of rest in 4 conditions: HL (%1-RM/cuff pressure: 80%/0); LL (20%/0); LBFR(sub DBP) (20%/1.3 x diastolic blood pressure, BP); LBFR(sub SBP) (20%/1.3 x supine systolic BP). The cuff remained inflated throughout the LBFR exercise sessions. Artery diameter, velocity time integral (VTI), and stroke volume (SV) were measured using Doppler ultrasound at rest and immediately after each set of exercise. Heart rate (HR) was monitored using a 3-lead ECG. BF was calculated as VTI x vessel cross-sectional area. CO was calculated as HR x SV. The data obtained after each set of exercise were averaged and used for analyses. Multi-level modeling was used to determine the effect of exercise condition on dependent variables. Statistical significance was set a priori at p< 0.05. RESULTS: Artery diameter did not change from baseline. BF increased (p<0.05) after exercise in each condition except LBFR(sub SBP) in the order of HL (12.73+/-1.42 cm3,mean+/-SE) > LL (9.92+/-0.82 cm3) > LBFR(sub dBP)(6.47+/-0.79 cm3) > LBFR(sub SBP) (3.51+/-0.59 cm3). Blunted exercise induced increases occurred in HR, SV, and CO after LBFR compared to HL and LL. HR increased 45% after HL and LL and 28% after LBFR (p<0.05), but SV increased (p<0.05) only after HL. Consequently, the increase (p<0.05) in CO was greater in HL and LL (approximately 3 L/min) than in LBFR (approximately 1 L/min). CONCLUSION: BF during LBFR(sub SBP) was 1/3 of that observed in LL, which supports the hypothesis that local ischemia stimulates the LBFR hypertrophic response. As the cuff did not compress the artery, the ischemia may have occurred because of the blunted rise in CO or because arterial BP cannot overcome the cuff pressure. As LBFR(sub DBP) effectively reduced BF and CO with cuff pressures less than systolic BP, future studies should investigate the hypertrophic potential of LBFR at even lower cuff pressures.

Everett, M. E.↗

NASA NTRS2013

The Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Mission Applications Study

The objective of the HIAD Mission Applications Study is to quantify the benefits of HIAD infusion to the concept of operations of high priority exploration missions. Results of the study will identify the range of mission concepts ideally suited to HIADs and provide mission-pull to associated technology development programs while further advancing operational concepts associated with HIAD technology. A summary of Year 1 modeling and analysis results is presented covering missions focusing on Earth and Mars-based applications. Recommended HIAD scales are presented for near term and future mission opportunities and the associated environments (heating and structural loads) are described.

Bose, David M.↗

NASA NTRS1982

Finite element analysis of wrinkling membranes

The finite element analysis of wrinkling membranes was investigated. The determination of stresses and deformations within large partly wrinkled membrane surfaces is a problem of significant technical interest in such areas as conceptual design and analysis of ultra lightweight spacecraft structures. A closed-form solution to an axisymmetric problem involving partial wrinkling of an inflated shallow membrane was obtained. In particular, a membrane in the shape of a sperical annulus was considered. The outer edge of the annulus was assumed to be fixed so that no displacements occur along the outer perimeter. The inner edge is assumed to be clamped to a rigid movable plug. Solutions for the complete stress, strain, and displacement fields under the assumption of inextensional material behavior are presented for the case of pure torsional loads applied to the plug, and for the case of pure axial loads applied to the plug.

Miller, R. K.↗