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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Stardust Entry Reconstruction

An overview of the reconstruction analyses performed for the Stardust capsule entry is described. The results indicate that the actual entry was very close to the pre-entry predictions. The capsule landed 8.1 km north-northwest of the desired target at Utah Test and Training Range. Analyses of infrared video footage and radar range data (obtained from tracking stations) during the descent show that drogue parachute deployment was 4.8 s later than the pre-entry prediction, while main parachute deployment was 19.3 s earlier than the pre-set timer indicating that main deployment was actually triggered by the backup baroswitch. Reconstruction of a best estimated trajectory revealed that the aerodynamic drag experienced by the capsule during hypersonic flight was within 1% of pre-entry predications. Observations of the heatshield support the pre-entry estimates of small hypersonic angles of attack, since there was very little, if any, charring of the shoulder region or the aftbody. Through this investigation, an overall assertion can be made that all the data gathered from the Stardust capsule entry were consistent with flight performance close to nominal pre-entry predictions. Consequently, the design principles and methodologies utilized for the flight dynamics, aerodynamics, and aerothermodynamics analyses have been corroborated.

Desai, Prasun N.↗
NASA NTRS1999

Entry Trajectory Issues for the Stardust Sample Return Capsule

The Stardust mission was successfully launched on February 7, 1999. It will be the first mission to return samples from a comet. The sample return capsule, which is passively controlled during the fastest Earth entry ever, will land by parachute in Utah. The present study describes the analysis of the entry, descent, and landing of the returning sample capsule utilizing the final, launch configuration capsule mass properties. The effects of two aerodynamic instabilities are revealed (one in the high altitude free molecular regime and the other in the transonic/subsonic flow regime). These instabilities could lead to unacceptably large excursions in the angle-of-attack near peak heating and main parachute deployment, respectively. To reduce the excursions resulting from the high altitude instability, the entry spin rate of the capsule is increased. To stabilize the excursions from the transonic/subsonic instability, a drogue chute with deployment triggered by a gravity-switch and timer is added prior to main parachute deployment. A Monte Carlo dispersion analysis of the modified entry (from which the impact of off-nominal conditions during the entry is ascertained) predicts that the capsule attitude excursions near peak heating and drogue chute deployment are within Stardust mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 60.8 km in downrange by 19.9 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗
NASA NTRS1999

Entry Trajectory Issues for the Stardust Sample Return Capsule

The Stardust mission was successfully launched on February 7, 1999. It will be the first mission to return samples from a comet. The sample return capsule, which is passively controlled during the fastest Earth entry ever, will land by parachute in Utah. The present study describes the analysis of the entry, descent, and landing of the returning sample capsule utilizing the final, launch configuration capsule mass properties. The effects of two aerodynamic instabilities are revealed (one in the high altitude free molecular regime and the other in the transonic/subsonic flow regime). These instabilities could lead to unacceptably large excursions in the angle-of-attack near peak heating and main parachute deployment, respectively. To reduce the excursions resulting from the high altitude instability, the entry spin rate of the capsule is increased. To stabilize the excursions from the transonic/subsonic instability, a drogue chute with deployment triggered by a gravity-switch and timer is added prior to main parachute deployment. A Monte Carlo dispersion analysis of the modified entry (from which the impact of off-nominal conditions during the entry is ascertained) predicts that the capsule attitude excursions near peak heating and drogue chute deployment are within Stardust mission limits. Additionally, the size of the resulting 3-sigma landing ellipse is 60.8 km in downrange by 19.9 km in crossrange, which is within the Utah Test and Training Range boundaries.

Desai, Prasun N.↗
NASA NTRS2005

Genesis Sample Return Capsule Overview

I. Simple Entry Capsule Concept: a) Spin-Stabilized/No Active Control Systems; b) Ballistic Entry for 11.04 km/sec Velocity; c) No Heatshield Separation During Entry; d) Parachute Deploy via g-Switch + Timer. II. Stardust Design Inheritance a) Forebody Shape; b) Seal Concepts; c) Parachute Deploy Control; d) Utah Landing Site (UTTR). III. TPS Systems a) Heatshield - Carbon-Carbon - First Planetary Entry; b) Backshell - SLA-561V - Flight Heritage from Pathfinder, MER; d) Forebody Structural Penetrations Aerothermal and TPS Design Process has the Same Methodology as Used for Pathfinder, MER Flight Vehicles.

Willcockson, Bill↗
NASA NTRS2011

Mars Science Laboratory Entry Guidance Improvements for Mars 2018 (DRAFT)

In 2011, the Mars Science Laboratory (MSL) will be launched in a mission to deliver the largest and most capable rover to date to the surface of Mars. A follow on MSL-derived mission, referred to as Mars 2018, is planned for 2018. Mars 2018 goals include performance enhancements of the Entry, Descent and Landing over that of its predecessor MSL mission of 2011. This paper will discuss the main elements of the modified 2018 EDL preliminary design that will increase performance on the entry phase of the mission. In particular, these elements will increase the parachute deploy altitude to allow for more time margin during the subsequent descent and landing phases and reduce the delivery ellipse size at parachute deploy through modifications in the entry reference trajectory design, guidance trigger logic design, and the effect of additional navigation hardware.

Garcia-Llama, Eduardo↗
NASA NTRS1973

An advanced technique for the prediction of decelerator system dynamics.

An advanced two-body six-degree-of-freedom computer model employing an indeterminate structures approach has been developed for the parachute deployment process. The program determines both vehicular and decelerator responses to aerodynamic and physical property inputs. A better insight into the dynamic processes that occur during parachute deployment has been developed. The model is of value in sensitivity studies to isolate important parameters that affect the vehicular response.

Talay, T. A.↗
NASA NTRS2008

A Study of the Effects of Atmospheric Phenomena on Mars Science Laboratory Entry Performance

At Earth during entry the shuttle has experienced what has come to be known as potholes in the sky or regions of the atmosphere where the density changes suddenly. Because of the small data set of atmospheric information where the Mars Science Laboratory (MSL) parachute deploys, the purpose of this study is to examine the effect similar atmospheric pothole characteristics, should they exist at Mars, would have on MSL entry performance. The study considers the sensitivity of entry design metrics, including altitude and range error at parachute deploy and propellant use, to pothole like density and wind phenomena.

Cianciolo, Alicia D.↗
NASA NTRS2023

Dragonfly: A Case Study in Dynamic Stability Characterization

Next decade, the Dragonfly probe will enter Titan’s atmosphere on a mission to explore and characterize its unique and complex organic chemistry. In contrast to Mars entry missions, where the challenge is to slow down in its tenuous atmosphere before impacting the surface, Titan’s thick atmosphere provides generous drag. However, Titan has its own unique EDL challenges. Among these challenges is the long time (on the order of 90 minutes!) the vehicle will need to descend unguided through flow regimes where blunt-body entry probes are known to be dynamically unstable. This elongated timeline, both prior to drogue parachute deployment, and during the ensuing descent, has the potential to allow adverse dynamic instability properties to produce large amplitude oscillations in angle-of-attack. This, in-turn may lead to non-compliance in vehicle attitude requirements for EDL events such as parachute deployment, and lander release, or in the most extreme case, tumbling of the aeroshell and near-assured loss of mission.

EDL↗
NASA NTRS2023

Dragonfly: A Case Study in Dynamic Stability Characterization

Next decade, the Dragonfly probe will enter Titan’s atmosphere on a mission to explore and characterize its unique and complex organic chemistry. In contrast to Mars entry missions, where the challenge is to slow down in its tenuous atmosphere before impacting the surface, Titan’s thick atmosphere provides generous drag. However, Titan has its own unique EDL challenges. Among these challenges is the long time (on the order of 90 minutes!) the vehicle will need to descend unguided through flow regimes where blunt-body entry probes are known to be dynamically unstable. This elongated timeline, both prior to drogue parachute deployment, and during the ensuing descent, has the potential to allow adverse dynamic instability properties to produce large amplitude oscillations in angle-of-attack. This, in-turn may lead to non-compliance in vehicle attitude requirements for EDL events such as parachute deployment, and lander release, or in the most extreme case, tumbling of the aeroshell and near-assured loss of mission.

EDL↗
NASA NTRS2013

GFEChutes Lo-Fi

NASA needed to provide a software model of a parachute system for a manned re-entry vehicle. NASA has parachute codes, e.g., the Descent Simulation System (DSS), that date back to the Apollo Program. Since the space shuttle did not rely on parachutes as its primary descent control mechanism, DSS has not been maintained or incorporated into modern simulation architectures such as Osiris and Antares, which are used for new mission simulations. GFEChutes Lo-Fi is an object-oriented implementation of conventional parachute codes designed for use in modern simulation environments. The GFE (Government Furnished Equipment), low-fidelity (Lo-Fi) parachute model (GFEChutes Lo-Fi) is a software package capable of modeling the effects of multiple parachutes, deployed concurrently and/or sequentially, on a vehicle during the subsonic phase of reentry into planetary atmosphere. The term "low-fidelity" distinguishes models that represent the parachutes as simple forces acting on the vehicle, as opposed to independent aerodynamic bodies. GFEChutes Lo-Fi was created from these existing models to be clean, modular, certified as NASA Class C software, and portable, or "plug and play." The GFE Lo-Fi Chutes Model provides basic modeling capability of a sequential series of parachute activities. Actions include deploying the parachute, changing the reefing on the parachute, and cutting away the parachute. Multiple chutes can be deployed at any given time, but all chutes in that case are assumed to behave as individually isolated chutes; there is no modeling of any interactions between deployed chutes. Drag characteristics of a deployed chute are based on a coefficient of drag, the face area of the chute, and the local dynamic pressure only. The orientation of the chute is approximately modeled for purposes of obtaining torques on the vehicle, but the dynamic state of the chute as a separate entity is not integrated - the treatment is simply an approximation. The innovation in GFEChutes Lo-Fi is to use an object design that closely followed the mechanical characteristics and structure of a physical system of parachutes and their deployment mechanisms. Software objects represent the components of the system, and use of an object hierarchy allows a progression from general component outlines to specific implementations. These extra chutes were not part of the baseline deceleration sequence of drogues and mains, but still had to be simulated. The major innovation in GFEChutes Lo-Fi is the software design and architecture.

Gist, Emily↗
NASA NTRS1973

Parachute mortar design.

Mortars are used as one method for ejecting parachutes into the airstream to decelerate spacecraft and aircraft pilot escape modules and to effect spin recovery of the aircraft. An approach to design of mortars in the class that can accommodate parachutes in the 20- to 55-foot-diameter size is presented. Parachute deployment considerations are discussed. Comments are made on the design of a power unit, mortar tube, cover, and sabot. Propellant selection and breech characteristics and size are discussed. A method of estimating hardware weights and reaction load is presented. In addition, some aspects of erodible orifices are given as well as comments concerning ambient effects on performance. This paper collates data and experience from design and flight qualification of four mortar systems, and provides pertinent estimations that should be of interest on programs considering parachute deployment.

Pleasants, J. E.↗
NASA NTRS1967

Performance of 26 Meter Diameter Ringsail Parachute in a Simulated Martian Environment

Inflation, drag, and stability characteristics of an 85.3-foot (26-meter) nominal diameter ringsail parachute deployed at a Mach number of 1.15 and at an altitude of 132,600 feet (40.42 kilometers) were obtained from the first flight test of the Planetary Entry Parachute Program. After deployment, the parachute inflated to the reefed condition. However, the canopy was unstable and produced low drag in the reefed condition. Upon disreefing and opening to full inflation, a slight instability in the canopy mouth was observed initially. After a short time, the fluctuations diminished and a stable configuration was attained. Results indicate a loss in drag during the fluctuation period prior to stable inflation. During descent, stability characteristics of the system were such that the average pitch-yaw angle from the local vertical was less than 10 degrees. Rolling motion between the payload and parachute canopy quickly damped to small amplitude.

Whitlock, Charles H.↗
NASA NTRS1968

Performance of a 16.6 Meter Diameter Modified Ringsail Parachute in a Simulated Martian Environment

Inflation, drag, and stability characteristics of a 54.5 -foot nominal-diameter (16.6-meter) modified ringsail parachute deployed in the wake of a 15-foot-diameter (4.6-meter) spacecraft traveling at a Mach number of 1.6 and a dynamic pressure equal to 11.6 psf (555 N/m(exp 2)) were obtained from the third balloon-launched flight test of the Planetary Entry Parachute Program. After deployment, the parachute inflated rapidly to a full condition, partially collapsed, and reinflated to a stable configuration. After reinflation, an average drag coefficient near 0.6 based on nominal surface area was obtained. During descent, an aerodynamic trim angle was observed in a plane near several torn sails. Amplitude of the trim was approximately 15 degrees and oscillation about trim was less than 11 degrees.

Whitlock, Charles H.↗
NASA NTRS1968

Performance of a 16.6 Meter Diameter Cross Parachute in a Simulated Martian Environment

Inflation and drag characteristics of a 54.4-foot (16.6 meter) nominal-diameter cross parachute, deployed at a Mach number of 1.65 and a dynamic pressure of 12.68 lb/sq f t (607.1 N/m(exp2)), were obtained from the fourth balloon-launched flight test of the Planetary Entry Parachute Program (PEPP). After deployment, the parachute quickly inflated to a full condition, partially collapsed, and then gradually reinflated while undergoing rapid oscillations between over-inflation and under-inflation. The oscillations began while the parachute was still at supersonic speeds and continued to low subsonic speeds well below an altitude of 90,000 feet (27.4 km). These canopy instabilities produced large cyclic variations in the parachute's drag coefficient. The average value of drag coefficient was about 0.8 to 0.9 at subsonic speeds and slightly lower at supersonic speeds. These drag coefficient values were based on the actual fabric surface area of the parachute canopy. The parachute sustained minor damage consisting of two canopy tears and abrasions and tears on the riser line. It is believed that this damage did not produce a significant change in the performance of the parachute.

Lundstrom, Reginald R.↗
NASA NTRSDate not supplied

Orion Artemis I Descent and Landing Performance

The Artemis I mission successfully demonstrated the Orion Multi-Purpose Crew Module ability to trigger parachute deployment, control angular rate and touch-down heading using reaction control jets while under parachutes, and automatically detect splashdown. Guidance and control logic is implemented to initiate para-chute events in a favorable dynamic state and select a touchdown heading direction that will minimize impact loads at splashdown. Detection of splashdown inhibits control to prevent firing reaction control jets under water and begin the transition into post-landing activities that cut parachutes and deploy the Crew Module Up-righting System (CMUS). This paper presents analysis of Orion descent and landing guidance and control performance for the Artemis I mission utilizing imagery taken from the vehicle, recorded flight data, and results from high fidelity simulations.

Artemis 1↗
NASA NTRS1985

Compact Antispin Parachute System

Costs and complexity of parachute system greatly reduced. Rigidtowline parachute system operates in three steps: (1) When aircraft begins to spin, parachute compartment opened up, and parachute unfolds; (2) Parachute deployed and spin terminated; and (3) Parachute released. Advantage of towline system allows use of much smaller parachute, reducing design loads on structural reinforcement. Also, complex pyrotechnic-deployment or jettison systems no longer necessary.

Whipple, R. D.↗
NASA NTRS1968

Performance of a 19.7 Meter Diameter Disk-Gap-Band Parachute in a Simulated Martian Environment

Inflation and drag characteristics of a 64.7-foot (19.7-meter) nominal-diameter disk-gap-band parachute deployed at a Mach number of 1.59 and a dynamic pressure of 11.6 psf (555 newtons per m(exp 2)) were obtained from the second balloon-launched flight test of the Planetary Entry Parachute Program. In addition, parachute stability characteristics during the subsonic descent portion of the test are presented. After deployment, the parachute rapidly inflated to a full condition, partially collapsed, and then reinflated to a stable configuration. After reinflation, an average drag coefficient of about 0.55 based on nominal surface area was obtained. The parachute exhibited good stability characteristics during descent. The only major damage to the parachute during the test was the tearing of two canopy panels; a loss of less than 0.5 percent of nominal surface area resulted.

TESTS↗
NASA NTRS2012

Determination of Barometric Altimeter Errors for the Orion Exploration Flight Test-1 Entry

The Exploration Flight Test 1 (EFT-1) mission is the unmanned flight test for the upcoming Multi-Purpose Crew Vehicle (MPCV). During entry, the EFT-1 vehicle will trigger several Landing and Recovery System (LRS) events, such as parachute deployment, based on on-board altitude information. The primary altitude source is the filtered navigation solution updated with GPS measurement data. The vehicle also has three barometric altimeters that will be used to measure atmospheric pressure during entry. In the event that GPS data is not available during entry, the altitude derived from the barometric altimeter pressure will be used to trigger chute deployment for the drogues and main parachutes. Therefore it is important to understand the impact of error sources on the pressure measured by the barometric altimeters and on the altitude derived from that pressure. The error sources for the barometric altimeters are not independent, and many error sources result in bias in a specific direction. Therefore conventional error budget methods could not be applied. Instead, high fidelity Monte-Carlo simulation was performed and error bounds were determined based on the results of this analysis. Aerodynamic errors were the largest single contributor to the error budget for the barometric altimeters. The large errors drove a change to the altitude trigger setpoint for FBC jettison deploy.

Brown, Denise L.↗