Search NASA⌕ Search

SEARCH · Search NASA

Results for “Supersonic Parachute”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent, and Landing of a Capsule Form Vehicle

This paper describes a critical portion of the work that has been done at NASA, Ames Research Center regarding the use of the commercially developed Dragon capsule as a delivery vehicle for the elements of a high priority Mars Sample Return mission. The objective of the investigation was to determine entry and landed mass capabilities that cover anticipated mission conditions. The "Red Dragon", Mars configuration, uses supersonic retro-propulsion, with no required parachute system, to perform Entry, Descent, and Landing (EDL) maneuvers. The propulsive system proposed for use is the same system that will perform an abort, if necessary, for a human rated version of the Dragon capsule. Standard trajectory analysis tools are applied to publically available information about Dragon and other legacy capsule forms in order to perform the investigation. Trajectory simulation parameters include entry velocity, flight path angle, lift to drag Ratio (L/D), landing site elevation, atmosphere density, and total entry mass, in addition engineering assumptions for the performance of the propulsion system are stated. Mass estimates for major elements of the overall proposed architecture are coupled to this EDL analysis to close the overall architecture. Three synodic launch opportunities, beginning with the 2022 opportunity, define the arrival conditions. Results state the relations between the analysis parameters as well as sensitivities to those parameters. The EDL performance envelope includes landing altitudes between 0 and -4 km referenced to the Mars Orbiter Laser Altimeter datum as well as minimum and maximum atmosphere density. Total entry masses between 7 and 10 mt are considered with architecture closure occurring between 9.0 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the Thermal Protection System (TPS) currently in use for Dragon missions shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass and grow allowance for a Mars Ascent Vehicle (MAV), Earth Return Vehicle (ERV), and mission unique equipment. The useful payload supports an architecture that receives a sample from another surface asset and sends it directly back to Earth for recovery in a high Earth orbit. The work shows that emerging commercial capabilities as well as previously studied EDL methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that will also use propulsive EDL techniques

EDL↗
NASA NTRSDate not supplied

Development of an Ejectable Data Recorder Ejection Mechanism for the Low Earth Orbit Flight Test of an Inflatable Decelerator

On November 10, 2022, the 1100kg (2,425 lbs.) LOFTID Reentry Vehicle (RV) was launched on a United Launch Alliance Atlas V as a secondary payload with the Joint Polar Surveyor System-2. The 6-meter diameter (~20 ft.) aeroshell (a type of heat shield) entered the atmosphere at 8 kilometers per second (18,000 miles per hour), and flew nominally, enduring the intended heat pulse that saw temperatures exceeding 1371˚C (2500˚F) on the front side while the payload skin remained only about 38˚C (100˚F). The RV exceeded Mach 30 and the heat-affected aeroshell withstood a pressure pulse that exerted 9g’s deceleration maintaining stable flight through the hypersonic, supersonic, transonic, and subsonic regimes to the parachute deployment. As part of the Agency’s strategic goal “to extend human presence deeper into space and to the moon for sustainable long-term exploration and utilization”, the LOFTID inflatable aerodynamic decelerator or aeroshell technology could one day help land humans on Mars. As with any flight test, data collection is of utmost importance. Without a data downlink and a possibility of the RV sinking before the recovery crew got to it, a secondary data collection method was introduced. The RV would eject a data recorder, which would have a duplicate copy of the on-board flight date, before splashdown and be retrieved separately. This paper discusses the development of the ejection mechanism used to eject the data recorder from the RV during the test flight. The development includes discussions of design constraints, a design overview, the testing program, and lessons learned throughout the process all the way through successful data recorder recovery.

mechanism↗
NASA NTRS2017

Estimates for the Aerodynamic Coefficients of Ringsail and Disk-Gap-Band Parachutes Operating on Mars

The creation of flight dynamics simulations is often part of the design, development, and flight operations of missions that include entry, descent, and landing (EDL) in their concept of operation. Such flight dynamics simulations include models for many of the EDL sub-systems. For EDL systems that include a parachute, models of the parachute’s aerodynamic characteristics must be created for inclusion in the flight dynamics simulations. Models are presented for the aerodynamic coefficients (i.e., C(D),C(Tot),C(T),C(N), and C(m,SLCP)) of Supersonic Ringsail (SSRS) and Disk-Gap-Band (DGB) parachutes as functions of total porosity, λ(T), Mach number, M, and total angle of attack, λ(T) (when necessary). The source aerodynamic coefficients data used for creating these models were obtained from a wind tunnel test of subscale parachutes. In this wind tunnel test, subscale parachutes of both parachute types were fabricated from two different fabrics with very different permeabilities. By varying the fabric permeability, while maintaining the parachute geometry constant, it was possible to vary λ(T). The fabric permeability test data necessary for the calculation of λ(T) were obtained from samples of the same fabrics used to fabricate the subscale parachutes. Although the models for the aerodynamic coefficients are simple polynomial functions of λ(T) and M, they are capable of producing good reproductions of the source data. The (λ(T), M) domains over which these models are applicable are clearly defined. The models and their domains are applicable to flight operations on Mars. The models presented in this report are not unique – other models could be created from the source data. To make the creation of alternative numerical models, the aerodynamic coefficients source data are included in this report.

Cruz, Juan R.↗
NASA NTRS2016

Construction of Supersonic Flight Dynamics Test (SFDT-2) Vehicle Monte Carlo Splashdown Footprints for Use in Range Safety and Recovery Operations

The Low-Density Supersonic Dececlerator (LDSD) project performed the second test flight of a supersonic inflatable device (SIAD) and ring-sail parachute on June 8, 2015 over the Pacific Ocean west of Kauai. In order for recovery ships to quickly extract the test vehicle hardware from the ocean, and in the interest of safety to the population of the islands of Kauai and Nihau, statistical estimates of the splashdown location had to be performed. This paper describes the modeling assumptions used to generate the Monti Carlo splashdown trajectory simulation results and the use of the splashdown probability ellipses in support of satisfying range safety requirements.

Strauss, William D.↗
NASA NTRS1961

Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3

A wind-tunnel investigation has been conducted to determine the parameters affecting the aerodynamic performance of drogue parachutes in the Mach number range from 1.6 to 3. Flow studies of both rigid and flexible-parachute models were made by means of high-speed schlieren motion pictures and drag coefficients of the flexible-parachute models were measured at simulated altitudes from about 50,000 to 120,000 feet. Porosity and Mach number were found to be the most important factors influencing the drag and stability of flexible porous parachutes. Such parachutes have a limited range of stable'operation at supersonic speeds, except for those with very high porosities, but the drag coefficient decreases rapidly with increasing porosity.

Maynard, Julian D.↗
NASA NTRS2012

Red Dragon: Low-cost Access to the Surface of Mars using Commercial Capabilities

We will discuss the feasibility of using a minimally-modified variant of a SpaceX Dragon capsule as a low-cost, large-capacity, near-term, Mars lander for scientific and human-precursor missions. We have been evaluating such a Red Dragon platform as an option for a Discovery Program mission concept. A Red Dragon lander has the potential to be low cost primarily because it would be derived from a routinely-flying spacecraft. Dragon is being developed to ferry cargo and crew to and from the International Space Station (ISS). The cargo variant is currently undergoing test flights, which will be followed by standard ISS cargo missions and, eventually, crewed flights. The human variant, unlike other Earth-return vehicles, appears to also have most of the capabilities necessary to land on Mars. In particular, it has a set of high-thrust, throttleable, storable bi-propellant Super- Draco engines integrated directly into the capsule which are intended for launch abort and powered landings on Earth. These thrusters suggest the possibility of a parachute-free, fully-propulsive deceleration at Mars from supersonic speeds to the surface. Concepts for large, human-relevant landers (see, e.g., [1]) also often employ supersonic retro-propulsion; Red Dragon's entry, descent, and landing approach would scale to those landers. Further, SpaceX's Falcon Heavy launch vehicle, currently under development and expected to have its first flight in 2013, will be capable of sending Dragon on a trajectory to Mars. We will discuss our motivation for exploring a Red Dragon lander, the primary technical questions which determine its feasibility, and the current results of our analysis. In particular, we will examine entry, descent, and landing (EDL) in detail. We will describe the modifications to Dragon necessary for interplanetary cruise, EDL, and operations on the Martian surface.

Karcz, John↗
NASA NTRS2013

Feasibility of a Dragon-Derived Mars Lander for Scientific and Human-Precursor Missions

A minimally-modified SpaceX Dragon capsule launched on a Falcon Heavy rocket presents the possibility of a new low-cost, high-capacity Mars lander for robotic missions. We have been evaluating such a "Red Dragon" platform as an option for the Icebreaker Discovery Program mission concept. Dragon is currently in service ferrying cargo to and from the International Space Station, and a crew transport version is in development. The upcoming version, unlike other Earth-return vehicles, exhibits most of the capabilities necessary to land on Mars. In particular, it has a set of high-thrust, throttleable, storable bi-propellant "SuperDraco" engines integrated directly into the capsule that are intended for launch abort and powered landings on Earth. These thrusters provide the possibility of a parachute-free, fully-propulsive deceleration at Mars from supersonic speeds to the surface, a descent approach which would also scale well to larger future human landers. We will discuss the motivations for exploring a Red Dragon lander, the current results of our analysis of its feasibility and capabilities, and the implications of the platform for the Icebreaker mission concept. In particular, we will examine entry, descent, and landing (EDL) in detail. We will also describe the modifications to Dragon necessary for interplanetary cruise, EDL, and operations on the Martian surface. Our analysis to date indicates that a Red Dragon lander is feasible and that it would be capable of delivering more than 1000 kg of payload to sites at elevations three kilometers below the Mars Orbiter Laser Altimeter (MOLA) reference, which includes sites throughout most of the northern plains and Hellas.

Karcz, John S.↗
NASA NTRS2013

Progress in Payload Separation Risk Mitigation for a Deployable Venus Heat Shield

A deployable decelerator known as the Adaptive Deployable Entry and Placement Technology (ADEPT) offers substantial science and mass savings for the Venus In Situ Explorer (VISE) mission. The lander and science payload must be separated from ADEPT during atmospheric entry. This paper presents a trade study of the separation system concept of operations and provides a conceptual design of the baseline: aft-separation with a subsonic parachute. Viability of the separation system depends on the vehicle's dynamic stability characteristics during deceleration from supersonic to subsonic speeds. A trajectory sensitivity study presented shows that pitch damping and Venusian winds drive stability prior to parachute deployment, while entry spin rate is not a driver of stability below Mach 5. Additionally, progress in free-flight CFD techniques capable of computing aerodynamic damping parameters is presented. Exploratory simulations of ADEPT at a constant speed of Mach number of 0.8 suggest the vehicle may have an oscillation limit cycle near 5 angle-of-attack. The proposed separation system conceptual design is thought to be viable.

Smith, Brandon P.↗
NASA NTRS2004

Low-Pressure Capability of NASA Glenn's 10- by 10-Foot Supersonic Wind Tunnel Expanded

Extremely low dynamic pressure Q conditions are desired for space-related research including the testing of parachute designs and other decelerator concepts for future vehicles landing on Mars. Therefore, the low-pressure operating capability of the Abe Silverstein 10- by 10-foot Supersonic Wind Tunnel (10 10 SWT) at NASA Glenn Research Center was recently increased. Successful checkout tests performed in the fall of 2002 showed significantly reduced minimum operating pressures in the wind tunnel.

Roeder, James W.↗
NASA NTRS1960

Preliminary Investigation of a Paraglider

A preliminary investigation of the aerodynamic and control characteristics of a flexible glider similar to a parachute in construction has been made at the Langley Research Center to evaluate its capabilities as a reentry glider. Preliminary weight estimates of the proposed vehicle indicate that such a structure can be made with extremely low wing loading. Maximum temperatures during the reentry maneuver might be held as low as about 1,500 F. The results of wind-tunnel and free-glide tests show that the glider when constructed of nonporous material performed extremely well at subsonic speeds and could be flown at angles of attack from about 200 to 900. At supersonic speeds the wing showed none of the unfavorable tendencies exhibited by conventional parachutes at these speeds, such as squidding and breathing. Several methods of packing and deploying the glider have been successfully demonstrated. The results of this study indicate that this flexible-lifting-surface concept may provide a lightweight controllable paraglider for manned space vehicles.

Rogallo, Francis M.↗
NASA NTRS2013

Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

Supersonic Inflatable Aerodynamic Decelerators (SI↗
NASA NTRS1973

Drag and stability characteristics of high-speed parachutes in the transonic range.

The results of wind tunnel tests of three parachute configurations in the wake of a cone-cylinder are presented. The tests were conducted to extend the drag and stability characteristics of selected parachutes through the transonic speed range. The configurations studied were the hemisflo ribbon, the cross, and the disk-gap-band types. The results are presented as the variation of the parachute drag coefficient with Mach number. General stability characteristics of the parachutes are discussed. The results are then correlated with some published subsonic and supersonic data.

Alexander, W. C.↗
NASA NTRS2018

Entry, Descent, and Landing Performance for a Mid-Lift-to-Drag Ratio Vehicle at Mars

In an effort to mature the design of the Mid-Lift-to-Drag ratio Rigid Vehicle (MRV) candidate of the NASA Evolvable Mars Campaign (EMC) architecture study, end-to-end six-degree-of-freedom (6DOF) simulations are needed to ensure a successful entry, descent, and landing (EDL) design. The EMC study is assessing different vehicle and mission architectures to determine which candidate would be best to deliver a 20 metric ton payload to the surface of Mars. Due to the large mass payload and the relatively low atmospheric density of Mars, all candidates of the EMC study propose to use Supersonic Retro-Propulsion (SRP) throughout the descent and landing phase, as opposed to parachutes, in order to decelerate to a subsonic touchdown. This paper presents a 6DOF entry-to-landing performance and controllability study with sensitivities to dispersions, particularly in the powered descent and landing phases.

mars↗
NASA NTRS2004

Inflatable Emergency Atmospheric-Entry Vehicles

In response to the loss of seven astronauts in the Space Shuttle Columbia disaster, large, lightweight, inflatable atmospheric- entry vehicles have been proposed as means of emergency descent and landing for persons who must abandon a spacecraft that is about to reenter the atmosphere and has been determined to be unable to land safely. Such a vehicle would act as an atmospheric decelerator at supersonic speed in the upper atmosphere, and a smaller, central astronaut pod could then separate at lower altitudes and parachute separately to Earth. Astronaut-rescue systems that have been considered previously have been massive, and the cost of designing them has exceeded the cost of fabrication of a space shuttle. In contrast, an inflatable emergency-landing vehicle according to the proposal would have a mass between 100 and 200 kg, could be stored in a volume of approximately 0.2 to 0.4 cu m, and could likely be designed and built much less expensively. When fully inflated, the escape vehicle behaves as a large balloon parachute, or ballute. Due to very low mass-per-surface area, a large radius, and a large coefficient of drag, ballutes decelerate at much higher altitudes and with much lower heating rates than the space shuttle. Although the space shuttle atmospheric reentry results in surface temperatures of about 1,600 C, ballutes can be designed for maximum temperatures below 600 C. This allows ballutes to be fabricated with lightweight ZYLON(Registered TradeMark) or polybenzoxazole (PBO), or equivalent.

Jones, Jack↗
NASA NTRS1974

Flight tests of Viking parachute system in three Mach number regimes. 1: Vehicle description, test operations, and performance

Flight qualifications for parachutes were tested on full-scale simulated Viking spacecraft at entry conditions for the Viking 1975 mission to Mars. The vehicle was carried to an altitude of 36.6 km for the supersonic and transonic tests by a 980.000 cu m balloon. The vehicles were released and propelled to test conditions with rocket engines. A 117,940 cu m balloon carried the test vehicle to an altitude of 27.5 km and the conditions for the subsonic tests were achieved in free fall. Aeroshell separation occurred on all test vehicles from 8 to 14 seconds after parachute deployment. This report describes: (1) the test vehicle; (2) methods used to insure that the test conditions were achieved; and (3) the balloon system design and operations. The report also presents the performance data from onboard and ground based instruments and the results from a statistical trajectory program which gives a continuous history of test-vehicle motions.

Lundstrom, R. R.↗
NASA NTRS2005

Enabling Pinpoint Landing on Mars

This slide presentation reviews the entry, descent and landing (EDL) of a pinpoint landing (PPL) on Mars. These PPL missions will be required to deliver about 1000 kg of useful payload to the surface of Mars, therefore soft landings are of primary interest. The landing sites will be in a mid to to high latitude with possible sites about 2.5 km above the martian mean surface altitude. The applicable EDL is described and reviewed in phases. The evaluation approach is reviewed and the requirements for an accurate landing are reviewed. The descent of the unguided aeroshell entry phase dispersion due to trajectory and ballistic coefficient variations are shown in charts. These charts view the dispersions from three entries, Entry from orbit, and two types of direct entry. There is discussion of the differences in steerable subsonic parachute control vs dispersions, and the propulsive phase delta velocity vs dispersions. Data is presented for the three trajectory phases (i.e., Aeroshell, supersonic and subsonic chute) for Direct entry and low orbit entry. The results of the analysis is presented, including possibilities for mitigation of dispersions. The analysis of the navigation error is summarized, and the trajectory biasing for martian winds is assessed.

aerodynamic↗
NASA NTRSDate not supplied

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↗