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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 163 records · Page 9

Nonapplicability of linear finite element programs to the stress analysis of tires

A static finite element stress analysis of an inflated radial car tire was carried out. The deformed shape of the sidewall presents outward bulging. The analysis of a homogeneous isotropic toroidal shell shows that the problem is common to all solids of this type. The study suggests that the geometric stiffness due to the inflation pressure has to be taken into account. Also, the resulting large displacements make it necessary for the geometry to be updated at each load step.

Durand, M.↗

TransHab Tapered Diamond Stitch Methods - Crewed Inflatable Softgoods Structures

The NASA TransHab project was a flight development program of a crewed inflatable habitat designed for in-space use. Through the course of the development, several novel technologies were developed and patented by NASA. One of these patents included information regarding a custom seam stitch that was used in the TransHab restraint layer design. As the patent protection has ended, the details of the custom seam stitch are being disclosed. The seam stitch developed during the TransHab program is known as the ‘tapered diamond stitch’ and includes both a ‘single’ and ‘double’ tapered diamond pattern for ‘end termination’ and ‘continuous’ webbing components respectively. It was designed for 1-in wide Kevlar webbings with load ratings of 6,000 lbf and 12,500 lbf. The stitch pattern was designed to gradually transfer load along the length of the seam using a tapered diamond geometry. The diamond stitch has been shown in tensile testing to provide greater than 85% load strength efficiency, compared to the pristine webbing.

transhab↗

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 Systems Study (DSS) testing campaign. A generalized, Adaptive Mesh Refinement (AMR) shock capturing and plume refinement technique was developed and calibrated for producing automatic and optimized grid systems unique to any given testing condition. Simulations were subjected to successive levels of grid refinement in order to determine solution independence from grid resolution. 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-sinsoidal expansions and contractions of the streamwise-oriented plume, resulting in unsteady aerodynamic loads on the vehicle.

Supersonic Retropropulsion↗

Flight Test of a 40-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 1.91 and a Dynamic Pressure of 11.6 Pounds per Square Foot

A 40-foot (12.2 meter) nominal-diameter disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator Program (SPED-I). The test parachute was ejected by a deployment mortar from an instrumented payload at an altitude of 140,000 feet (42.5 kilometers). The payload was at a Mach number of 1.91 and the dynamic pressure was 11.6 pounds per square foot (555 newtons per square meter) at the time the parachute deployment mortar was fired. The parachute reached suspension line stretch in 0.43 second with a resultant snatch force loading of 1990 pounds (8850 newtons). The maximum parachute opening load of 6500 pounds (28,910 newtons) came 0.61 second later at a total elapsed time from mortar firing of 1.04 seconds. The first full inflation occurred at 1.12 seconds and stable inflation was achieved at approximately 1.60 seconds. The parachute had an average axial-force coefficient of 0.53 during the deceleration period. During the steady-state descent portion of the flight test, the average effective drag coefficient was also 0.53 and pitch-yaw oscillations of the canopy averaged less than 10 degrees in the altitude region above 100,000 feet (30.5 meters).

EXPERIMENTS↗

Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Mission Overview, Science Return, and Future Applications of This Technology

The Low-Earth Orbit (LEO) Flight Test of an Inflatable Decelerator (LOFTID) mission was the culmination of two decades of research and development for Hyper-sonic Inflatable Aerodynamic Decelerator (HIAD) technology. LOFTID was a project overseen by the Technology Demonstration Mission (TDM) program within NASA’s Space Technology Mission Directorate. The success of the LOFTID mission could enable new NASA missions to Mars, Venus, and most solar-system destinations with atmospheres, as well as cost-effective payload returns to Earth, including in-space manufactured materials and launch vehicle asset recovery. LOFTID, with its unique inflatable heat shield, was the first-of-a-kind orbital reentry flight, and the largest blunt body atmospheric entry of any kind. On November 10, 2022, just over 10 years since the previous flight test of the smaller sub-orbital Inflatable Reentry Vehicle Experiment-3 (IRVE-3) [1], NASA Langley Research Center, with partner United Launch Alliance (ULA), successfully launched and achieved reentry and recovery of the LOFTID Reentry Vehicle (RV), further demonstrating the viability of the HIAD technology for large-diameter, inflatable heat shields to safely and accurately deliver large payloads through an atmosphere via a controlled descent and landing. Launching as a secondary payload with the Joint Polar Satellite System 2 (JPSS-2) from Vandenberg Space Force Base, California, stowed inside an ex-tended payload adapter of the Atlas V 401 launch vehicle, the LOFTID mission officially began after the JPSS-2 payload was delivered to its orbit by the Centaur second stage. The LOFTID RV was flying solo on its spin-stabilized ballistic reentry trajectory about one hour after launch, and the flight ended approximately one hour later with a gentle splashdown under parachute in the Pacific Ocean off the east coast of Hawaii, where the RV was recovered and later shipped back to NASA Langley. LOFTID endured the harsh environments of atmospheric reentry while demonstrating stable aerodynamics through the entire spectrum of hypersonic, supersonic, transonic, and subsonic flight. The LOFTID RV was exposed to an aeroheating environment representative of many Mars and LEO HIAD applications, while successfully demonstrating the ability of the heat-affected inflatable structure to with-stand aerodynamic forces that exceeded those expected at Mars. This flight demonstration of a 6m diameter HIAD confirmed the technology structural and thermal performance as it protected the 1100kg RV entering Earth’s atmosphere at 8 km/s, reaching Mach 30, and experiencing 9g deceleration before deploying parachutes and splashing down in the Pacific Ocean. HIAD technology involves an aeroshell that can be hard packed into a small volume for launch and then deployed prior to atmospheric entry for a controlled deceleration through the atmosphere. Large deployable heat shields enable spacecraft to carry bigger, heavier payloads, including scientific instruments and human support systems for planetary landing and exploration. Much larger than traditional fixed diameter aeroshells that are constrained by the size of launch vehicle shrouds, inflatable decelerators create more drag and start the deceleration process in the upper reaches of the atmosphere with greater efficiency and stability. LOFTID’s successful demonstration of the HIAD technology has greatly expanded NASA’s options for future planetary missions and opened new commercial opportunities for lower-cost mission pay-load recovery including recovery of launch vehicle assets, or as a delivery system to return cargo from the International Space Station or cislunar space. The developments for some of these applications are al-ready underway. While LOFTID was indeed a first-of-a-kind flight for an inflatable heatshield, the largest blunt-body entry ever, its remarkable performance assured that it will not be the last of its kind.

John DiNonno↗

A Numerical Investigation of Parachute Deployment in Supersonic Flow

A validated computational fluid-structure interaction method applied to supersonic parachute inflation is extended to consider a unique, parallel self-contact algorithm, porosity interface conditions on the canopy, improved flow field sampling procedures for obtaining quality loading on the parachute canopy, adaptive mesh refinement, and improved treatment of thin geometries in an immersed boundary framework. These extensions are discussed in detail and demonstrated individually on test problems. Finally, the developments are brought together for demonstration on a sub-scale MSL parachute geometry.

Boustani, Jonathan↗

Photogrammetry of a Hypersonic Inflatable Aerodynamic Decelerator

In 2012, two large-scale models of a Hypersonic Inflatable Aerodynamic decelerator were tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. One of the objectives of this test was to measure model deflections under aerodynamic loading that approximated expected flight conditions. The measurements were acquired using stereo photogrammetry. Four pairs of stereo cameras were mounted inside the NFAC test section, each imaging a particular section of the HIAD. The views were then stitched together post-test to create a surface deformation profile. The data from the photogram- metry system will largely be used for comparisons to and refinement of Fluid Structure Interaction models. This paper describes how a commercial photogrammetry system was adapted to make the measurements and presents some preliminary results.

aerodynamic brakes↗

Aerial Deployment and Inflation System for Mars Helium Balloons

A method is examined for safely deploying and inflating helium balloons for missions at Mars. The key for making it possible to deploy balloons that are light enough to be buoyant in the thin, Martian atmosphere is to mitigate the transient forces on the balloon that might tear it. A fully inflated Mars balloon has a diameter of 10 m, so it must be folded up for the trip to Mars, unfolded upon arrival, and then inflated with helium gas in the atmosphere. Safe entry into the Martian atmosphere requires the use of an aeroshell vehicle, which protects against severe heating and pressure loads associated with the hypersonic entry flight. Drag decelerates the aeroshell to supersonic speeds, then two parachutes deploy to slow the vehicle down to the needed safe speed of 25 to 35 m/s for balloon deployment. The parachute system descent dynamic pressure must be approximately 5 Pa or lower at an altitude of 4 km or more above the surface.

Lachenmeler, Tim↗

Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 2.72 and a Dynamic Pressure of 9.7 Pounds per Square Foot

A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator (SPED-I) Program. The test parachute was deployed from an instrumented payload by means of a deployment mortar when the payload was at an altitude of 158,500 feet (48.2 kilometers), a Mach number of 2.72, and a free-stream dynamic pressure of 9.7 pounds per foot(exp 2) (465 newtons per meter(exp 2)). Suspension line stretch occurred 0.46 second after mortar firing and the resulting snatch force loading was -8.lg. The maximum acceleration experienced by the payload due to parachute opening was -27.2g at 0.50 second after the snatch force peak for a total elapsed time from mortar firing of 0.96 second. Canopy-shape variations occurred during the higher Mach number portion of the flight test (M greater than 1.4) and the payload was subjected to large amplitude oscillatory loads. A calculated average nominal axial-force coefficient ranged from about 0.25 immediately after the first canopy opening to about 0.50 as the canopy attained a steady inflated shape. One gore of the test parachute was damaged when the deployment bag with mortar lid passed through it from behind approximately 2 seconds after deployment was initiated. Although the canopy damage caused by the deployment bag penetration had no apparent effect on the functional capability of the test parachute, it may have affected parachute performance since the average effective drag coefficient of 0.48 was 9 percent less than that of a previously tested parachute of the same configuration.

EXPERIMENTS↗

Non-Axisymmetric Inflatable Pressure Structure (NAIPS) Concept that Enables Mass Efficient Packageable Pressure Vessels with Sealable Openings

Achieving minimal launch volume and mass are always important for space missions, especially for deep space manned missions where the costs required to transport mass to the destination are high and volume in the payload shroud is limited. Pressure vessels are used for many purposes in space missions including habitats, airlocks, and tank farms for fuel or processed resources. A lucrative approach to minimize launch volume is to construct the pressure vessels from soft goods so that they can be compactly packaged for launch and then inflated en route or at the final destination. In addition, there is the potential to reduce system mass because the packaged pressure vessels are inherently robust to launch loads and do not need to be modified from their in-service configuration to survive the launch environment. A novel concept is presented herein, in which sealable openings or hatches into the pressure vessels can also be fabricated from soft goods. To accomplish this, the structural shape is designed to have large regions where one principal stress is near zero. The pressure vessel is also required to have an elongated geometry for applications such as airlocks.

Doggett, William R.↗

Flexible Ablators: Applications and Arcjet Testing

Flexible ablators were conceived in 2009 to meet the technology pull for large, human Mars Exploration Class, 23 m diameter hypersonic inflatable aerodynamic decelerators. As described elsewhere, they have been recently undergoing initial technical readiness (TRL) advancement by NASA. The performance limits of flexible ablators in terms of maximum heat rates, pressure and shear remain to be defined. Further, it is hoped that this emerging technology will vastly expand the capability of future NASA missions involving atmospheric entry systems. This paper considers four topics of relevance to flexible ablators: (1) Their potential applications to near/far term human and robotic missions (2) Brief consideration of the balance between heat shield diameter, flexible ablator performance limits, entry vehicle controllability and aft-body shear layer impingement of interest to designers of very large entry vehicles, (3) The approach for developing bonding processes of flexible ablators for use on rigid entry bodies and (4) Design of large arcjet test articles that will enable the testing of flexible ablators in flight-like, combined environments (heat flux, pressure, shear and structural tensile loading). Based on a review of thermal protection system performance requirements for future entry vehicles, it is concluded that flexible ablators have broad applications to conventional, rigid entry body systems and are enabling to large deployable (both inflatable and mechanical) heat shields. Because of the game-changing nature of flexible ablators, it appears that NASA's Office of the Chief Technologist (OCT) will fund a focused, 3-year TRL advancement of the new materials capable of performance in heat fluxes in the range of 200-600 W/sq. cm. This support will enable the manufacture and use of the large-scale arcjet test designs that will be a key element of this OCT funded activity.

Thermal Protection↗

Development of 2 underseat energy absorbers for application to crashworthy passenger seats for general aviation aircraft

This report presents the methodology and results of a program conducted to develop two underseat energy absorber (E/A) concepts for application to nonadjustable crashworthy passenger seats for general aviation aircraft. One concept utilizes an inflated air bag, and the other, a convoluted sheet metal bellows. Prototypes of both were designed, built, and tested. Both concepts demonstrated the necessary features of an energy absorber (load-limiter); however, the air bag concept is particularly encouraging because of its light weight. Several seat frame concepts also were investigated as a means of resisting longitudinal and lateral loads and of guiding the primary vertical stroke of the underseat energy absorber. Further development of a seat system design using the underseat energy absorbers is recommended because they provide greatly enhanced crash survivability as compared with existing general aviation aircraft seats.

Warrick, J. C.↗

Passive Thermal Control of the Primordial Inflation Polarization Explorer (PIPER) Flight Electronics

The scientific balloon program at NASA offers an exciting and open area of opportunity for testing new technologies and for conducting meaningful experimentation at a fraction of the cost of a space mission. This paper outlines a simple thermal model developed and employed for the Primordial Inflation Polarization ExplorER (PIPER). The sub-orbital environment that PIPER operates in hosts an interesting mix of atmospheric and space thermal challenges. The work done was to mitigate thermal loads and passively cool the payload's exterior mounted electronics at an altitude of 36.6 km. This was done by characterizing the thermal environment and then designing solutions for the heat loads through a combined radiation and conduction passive cooling radiator system thermal model. Despite the simplicity and subsequent limitations of the model, as well as some unexpected payload operational events, the model produced results between 0.31% and 11.8% difference between the predicted values and measured average temperatures. From these results, the model was able to successfully provide estimates for the electronics temperatures. Additional flights will be needed to eliminate unknowns encountered in this flight in order to further refine the model for higher accuracy.

Walts, Alexander R.↗

Creep Testing of Vectran Yarn

NASA and industry are developing inflatable space structures to provide deployable habitable volumes for future exploration missions in space and on the surface of the Moon and Mars. These structures must withstand the constant pressure loading of multi-year missions. These missions therefore require excellent creep resistance in the constituent cordage and webbings that provide the primary structure of the vessel, which is typically constructed using high-strength synthetic materials such as Vectran. Real-time and accelerated creep testing of Vectran webbings has been performed at NASA Langley Research Center (LaRC), however, acceptable correlation between the results of the two approaches has not yet been achieved. Given the additional complexity of the webbing architecture, creep testing at the yarn level is being studied to reduce the number of interacting properties and focus on the core creep behavior to see if better correlation can be achieved between real-time and accelerated tests. To achieve this objective a creep test stand that is able to support simultaneous testing of five Vectran yarns had been previously designed and built. The test stand utilizes two clamp-wrap grips to mount the yarn and has a 10:1 load ratio, via a lever-arm, henceforth, known as the lever-arm test stand. Real-time creep testing (without measuring the strain of the yarn) in this fixture produces yarn ruptures in the gauge section. However, a large scatter in the times to failure was seen in the results across multiple loading conditions. For example, a test at 70% of the ultimate tensile strength (UTS) was found to have specimens that failed within a matter of hours and others that took over 6 weeks to fail. Prior test data on Vectran webbings had indicated high variability in the material strength, leading to higher than average scatter in the creep testing of webbings, which could translate to the scatter in creep behavior at the yarn level as well. However, there was still the possibility that the lever-arm test stand was causing the high scatter in data; specifically, due to potential variability in the loading of each specimen. Therefore, an additional creep test setup was designed and built to attempt to determine if the large scatter in the results was due to inherent material properties, the lever-arm test stand, or both. The newly designed creep test stand, henceforth known as the conventional test stand, utilized the more typical method for creep testing which involved hanging a weight from the yarn with no additional load amplification. The conventional test stand was designed and fabricated in-house at LaRC. High speed cameras were used to validate that the specimens in the conventional test stand failed in the gauge section. Once it was determined that the conventional test stand would provide valid failures in the gauge section, creep tests on the Vectran yarns were run to compare to the lever-arm test stand results. Preliminary results for a 70% UTS creep test have some specimens failing within a number of hours and some failing in excess of 6 weeks. These results are similar to those seen on the lever-arm test stand, pointing towards a large variability inherent in the material. The poster/presentation will focus on the design, fabrication, and validation of the conventional test stand as well as present preliminary results comparing the creep data from the conventional and lever-arm test stands

creep testing↗

Development of a Compact, Low Cost Test Fixture to Evaluate Creep in High Strength Softgoods Materials under Constant Environmental Control

Space exploration is typically driven by two key factors: cost and payload mass. Therefore creating relatively low cost, low mass space hardware is critical. With the advancement of new high-strength flexible materials, a new chapter is opening in the evolution of pressurized space structures. Space-rated inflatable structures are attractive due to their low mass, compact stowage, and capability of performing as well, or better than, their rigid counterparts for applications such as habitats and airlocks. The lifetime properties of materials used for inflatable structures are not as well characterized as heritage composite and metallic materials; therefore, they require extensive testing to increase the level of confidence in their use. Creep testing is one method of evaluating the expected lifetime of materials that may be under constant load for many years in operational use. Conventional creep testing of high-strength softgoods webbings or cordage requires a large, temperature-controlled facility, with structural frames capable of supporting large dead weights, weighing thousands of pounds on each specimen. This is expensive and takes up a significant amount of volume in a facility for a long time. This paper describes the design and analysis of a low-cost, low-volume creep test stand that can be locally environmentally controlled, so that smaller, cheaper facilities can be used for characterizing the creep life of the softgoods materials used in inflatable space structures.

Softgoods↗

Low Velocity Airdrop Tests of an X-38 Backup Parachute Design

The NASA Johnson Space Center's X-38 program designed a new backup parachute system to recover the 25,000 lb X-38 prototype for the Crew Return Vehicle spacecraft. Due to weight and cost constraints, the main backup parachute design incorporated rapid and low cost fabrication techniques using off-the-shelf materials. Near the vent, the canopy was constructed of continuous ribbons, to provide more damage tolerance. The remainder of the canopy was a constructed with a continuous ringslot design. After cancellation of the X-38 program, the parachute design was resized, built, and drop tested for Natick Soldiers Center's Low Velocity Air Drop (LVAD) program to deliver cargo loads up to 22,000 lbs from altitudes as low as 500 feet above the ground. Drop tests results showed that the 500-foot LVAD parachute deployment conditions cause severe skirt inversion and inflation problems for large parachutes. The bag strip occurred at a high angle of attack, causing skirt inversion before the parachute could inflate. The addition of a short reefing line prevented the skirt inversion. Using a lower porosity in the vent area, than is normally used in large parachutes, improved inflation. The drop testing demonstrated that the parachute design could be refined to meet the requirements for the 500-foot LVAD mission.

Stein, Jenny M.↗

Fluid Structure Interaction of Parachutes in Supersonic Planetary Entry

A research program to provide physical insight into disk-gap-band parachute operation in the supersonic regime on Mars was conducted. The program included supersonic wind tunnel tests, computational fluid dynamics and fluid structure interaction simulations. Specifically, the nature and cause of the "area oscillation" phenomenon were investigated to determine the scale, aerodynamic, and aero-elastic dependence of the supersonic parachute collapse and re-inflation event. A variety of non-intrusive, temporally resolved, and high resolution diagnostic techniques were used to interrogate the flow and generate validation datasets. The results of flow visualization, particle image velocimetry, load measurements, and photogrammetric reconstruction will be presented. Implications to parachute design, use, and verification will also be discussed.

parachute↗

Hybrid Inflatable Pressure Vessel

Figure 1 shows a prototype of a large pressure vessel under development for eventual use as a habitable module for long spaceflight (e.g., for transporting humans to Mars). The vessel is a hybrid that comprises an inflatable shell attached to a rigid central structural core. The inflatable shell is, itself, a hybrid that comprises (1) a pressure bladder restrained against expansion by (2) a web of straps made from high-strength polymeric fabrics. On Earth, pressure vessels like this could be used, for example, as portable habitats that could be set up quickly in remote locations, portable hyperbaric chambers for treatment of decompression sickness, or flotation devices for offshore platforms. In addition, some aspects of the design of the fabric straps could be adapted to such other items as lifting straps, parachute straps, and automotive safety belts. Figure 2 depicts selected aspects of the design of a vessel of this type with a toroidal configuration. The bladder serves as an impermeable layer to keep air within the pressure vessel and, for this purpose, is sealed to the central structural core. The web includes longitudinal and circumferential straps. To help maintain the proper shape upon inflation after storage, longitudinal and circumferential straps are indexed together at several of their intersections. Because the web is not required to provide a pressure seal and the bladder is not required to sustain structural loads, the bladder and the web can be optimized for their respective functions. Thus, the bladder can be sealed directly to the rigid core without having to include the web in the seal substructure, and the web can be designed for strength. The ends of the longitudinal straps are attached to the ends of the rigid structural core by means of clevises. Each clevis pin is surrounded by a roller, around which a longitudinal strap is wrapped to form a lap seam with itself. The roller is of a large diameter chosen to reduce bending of the fibers in the strap. The roller also serves to equalize the load in the portions of the strap on both sides of the clevis pin. The lap seam is formed near the clevis by use of a tapered diamond stitch: This stitch is designed specifically to allow fibers in the stitch and strap to relax under load in such a manner that the load becomes more nearly evenly distributed among all fibers in the stitch region. Thus, the tapered diamond stitch prevents load concentrations that could cause premature failure of the strap and thereby increases the strength of the strap/structural-core joint. The lap seam can be rated at >90 percent of the strength of the strap material.

Raboin, Jasen↗