Results for “backshell” — Page 4

NASA NTRS2021

Overview of Dragonfly Entry Aerosciences Measurements (DrEAM)

NASA Ames Research Center (ARC) leads the Dragonfly Entry Aerosciences Measurements (DrEAM) project, which is an aeroshell instrumentation suite on the Dragonfly mission that fulfills the Engineering Science Investigation requirement for the New Frontiers mission. NASA ARC is partnering with NASA Langley Research Center (LaRC) and the German Aerospace Center (DLR) to provide a comprehensive sensor suite, including a DLR-provided Data Acquisition System (DAS). DrEAM will provide key aerothermodynamic data and performance analysis for Dragonfly’s forebody and backshell Thermal Protection System (TPS). Titan’s atmosphere predominantly consists of ni-trogen (~98% by mole) with small amounts of me-thane (~2% by mole) and other trace gases. CN is a strong radiator, and is found in nonequilibrium con-centrations for Titan entry. The accurate modeling of nonequilibrium CN radiation has proven to be a diffi-cult task. Prompted by the Huygens mission, many experimental campaigns and analyses were performed to better understand the aerothermal environments experienced by the probe during Titan entry [1]. How-ever, the Huygens probe carried no heatshield instru-mentation. Therefore, the DrEAM sensor suite will sig-nificantly advance the state-of-the-art not only by documenting the environment and performance of Dragonfly’s entry system but also by making key in situ measurements in Titan’s atmosphere for the first time. Aerothermal environments and TPS response will be measured using sensors whose flight heritage is tak-en from the Mars Entry, Descent, and Landing In-strumentation 2 (MEDLI2) thermocouple plugs and the COMbined Aerothermal and Radiometer Sensor (COMARS) suite [2], with the latter supplied by DLR. The MEDLI2 project used embedded thermocouples to directly measure the in-depth TPS temperature-time history at several locations on the heat shield and backshell of the Mars 2020 entry vehicle. These tem-perature measurements, in turn, can be used to infer surface environments via an inverse analysis proce-dure analogous to that used for MEDLI. For DrEAM, the thermocouple plug subsystem will be known as Dragonfly Sensors for Aero-Thermal Reconstruction (DragSTR). On Schiaparelli, the COMARS suite in-cluded three total surface-mounted heat flux sensors, three pressure sensors, and one radiometer. For DrEAM, the COMARS package will be known as the COmbined Sensor System for Titan Atmosphere (COSSTA). Since the methane concentration in the Titan atmos-phere is directly proportional to the radiative heat flux, the COSSTA measurements will be used to reduce the current uncertainty in the methane volume fraction. Atmospheric density measurements and capsule aero-dynamic data will be obtained through the onboard Inertial Measurement Unit (IMU), supplemented by pressure transducers similar to those used by the MEDLI and MEDLI2 projects. The DrEAM pressure sensors will be known as the Dragonfly Atmospheric Flight Transducers. (DrAFT). The pressure measure-ments, when combined with data from the on-board IMU, will allow for reconstruction of such quantities as vehicle Mach number, freestream density, and atmos-pheric winds. DrAFT measurements will enhance Dragonfly trajectory reconstruction and enable a sepa-ration of the aerodynamics from the atmosphere, as was done for MEDLI [3] and is currently in process for the MEDLI2 flight data set.

EDL↗

NASA NTRSDate not supplied

Inverse Estimation of Mars 2020 Entry Aeroheating Environments Using MEDLI2 Flight Data

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite collected heating and pressure data during entry into Mars' atmosphere of the Mars 2020 Perseverance rover. MEDLI2 included thermocouples, heat flux sensors, and pressure transducers on both the heatshield and the backshell. This paper covers the inverse estimation of heatshield and backshell surface heating based on the MEDLI2 Instrumented Sensor Plugs (MISPs), a network of thermocouples embedded in thermal protection system plugs across the aeroshell. Monte Carlo analysis was conducted to assess the sensitivity of the surface heat rate and temperature to uncertainties in thermocouple depth and material properties such as density, specific heat capacity, and thermal conductivity. Data from each MISP was also used to estimate the local time of transition from laminar to turbulent flow at each plug location.

Hannah S Alpert↗

NASA NTRS2022

Mars Entry Instrumentation Flight Data and Mars 2020 Entry Environments

On February 18th, 2021, the Mars 2020 entry vehicle delivered the Perseverance rover to the surface of Mars. The entry vehicle carried a set of instrumentation installed on the heatshield and backshell to measure aerodynamic and aerothermal performance, named the Mars Entry, Descent, and Landing Instrumentation 2. This set of instrumentation included pressure transducers, thermocouples, heatflux sensors, and a radiometer, as well as a dedicated sensor support electronics system. All MEDLI2 hardware operated as expected during cruise and entry. MEDLI2 sensors gathered accurate pressure measurements in hypersonic through supersonic regimes to reconstruct vehicle attitude and atmospheric profiles. MEDLI2 on the heatshield sensors indicated that surface temperatures, caused by turbulent heating beginning 70 seconds after entry, remained at or below 1430 °C, while heatshield bondline temperatures rose less than 45 °C. Backshell surface TPS temperatures peaked at 630 °C, which was caused primarily by radiative heating measured by several separate sensors. The MEDLI2 temperature and pressure measurements enabled further detailed characterization of the Mars 2020 entry performance, and the flight dataset will provide a wealth of information for the EDL community and future mission designers.

thermal protection system↗

NASA NTRSDate not supplied

Mars Entry Instrumentation Flight Data and Mars 2020 Entry Environments

On February 18th, 2021, the Mars 2020 entry vehicle delivered the Perseverance rover to the surface of Mars. The entry vehicle carried a set of instrumentation installed on the heatshield and backshell to measure aerodynamic and aerothermal performance, named the Mars Entry, Descent, and Landing Instrumentation 2. This set of instrumentation included pressure transducers, thermocouples, heatflux sensors, and a radiometer, as well as a dedicated sensor support electronics system. All MEDLI2 hardware operated as expected during cruise and entry. MEDLI2 sensors gathered accurate pressure measurements in hypersonic through supersonic regimes to reconstruct vehicle attitude and atmospheric profiles. MEDLI2 on the heatshield sensors indicated that surface temperatures, caused by turbulent heating beginning 70 seconds after entry, remained at or below 1430 °C, while heatshield bondline temperatures rose less than 45 °C. Backshell surface TPS temperatures peaked at 630 °C, which was caused primarily by radiative heating measured by several separate sensors. The MEDLI2 temperature and pressure measurements enabled further detailed characterization of the Mars 2020 entry performance, and the flight dataset will provide a wealth of information for the EDL community and future mission designers.

thermal protection systems↗

NASA NTRSDate not supplied

Recent Results from Dragonfly Testing/Analysis as we head to PDR

Dragonfly is a relocatable lander mission to Saturn's moon Titan4, which as well as being a target of out-standing astrobiological interest as an organic-rich Ocean World, has the combination of low gravity (1/7 that of Earth) and a thick atmosphere (4x the density of Earth), making it an environment uniquely suitable for flight. Thus, the Dragonfly lander (similar in size to the Curiosity Mars rover) can take off using lift from a set of eight rotors and fly to a new landing site several kilometers away. The ability to perform such flights, lasting approximately 30 minutes, every month or so on Titan brings unprecedented mobility to planetary exploration, on a world known to have a diverse land-scape of dunes, craters and other features. Dragonfly is planned to launch in 2027, and following a nearly seven year interplanetary cruise would arrive at Titan by 2034. Due to the large scale height of the Titan atmosphere, Entry, Descent, and Landing (EDL) will be prolonged affair, taking nearly two hours to reach the surface. The ballistic entry environments that Dragonfly will be subjected to are fairly similar to that experienced by recent Mars missions; peak heating on the aeroshell will be about 300 W/sq.cm and peak deceleration is about 10g’s. Following the five minute entry segment, much of the remaining time is spent descending on the drogue and main para-chutes, which carry the dual role of decelerating the spacecraft and stabilizing the system during the long descent. While on one hand, this leisurely EDL sequence affords a relaxed timeline and plenty of time for event staging, it also provides ample opportunity for small disturbances to grow into potential flight safety risks, adding emphasis to the need for careful modeling, simulation and testing of key dynamic events. About two hours after entering the atmosphere, the nearly one metric ton rotocraft will be lowered approximately one meter out of the backshell (the ‘pose’ maneuver) to expose all eight rotors. The rotors will then be used to arrest any residual spin rate and prepare the system for transition to powered flight. Once despin is complete and the lander reaches a target altitude of 1.2 km above the surface (as verified by on-board lidar), the lander will be released and free fall for approximately one second before beginning controlled free flight. This entire “preparation for powered flight” process takes place over several minutes while the system is subject to the dynamic environment produced by so-called “wrist-mode” oscillations as the lander and backshell swing on the main parachute. Once in free flight, the lander will engage on-board terrain relative navigation to locate and navigate to a safe landing zone in the Shangri-La dune field south of Selk crater. Communication during this sequence will be limited to a series of direct-to-Earth X-band tones signalling key events and providing forensic information. Once on the ground, the lander will begin to send additional information, including data collected during this EDL sequence by the on-board Dragonfly Entry Aerosciences Measurements (DrEAM) instrumentation suite. This presentation will walk through the entry to first landing timeline in more detail, with a focus on recent analysis and testing results that inform system performance, margins and residual risk estimation.

Dragonfly↗

NASA NTRSDate not supplied

Exploring the Accuracy of RANS Simulations for Mars Entry Vehicles

Accurate yet inexpensive predictions of aerodynamic coefficients for Mars entry vehicles have remained a consistent challenge over the past five decades. Below Mach 6, drag on the backshell becomes significant and must be accurately predicted. Steady Reynolds-averaged Navier-Stokes (RANS) models are commonly used, despite their poor predictions of backshell pressure. While scale-resolving simulations have shown promise in the past decade, there is still a need for cheap, accurate RANS predictions for large aerodynamic databases. The Mars Science Laboratory (MSL) is used as a case study to examine predictive accuracy and known shortcomings for RANS predictions of Mars entry vehicles. Several different grid generation techniques are compared, including a comparison between prismatic boundary layer grids and fully unstructured, tetrahedral grids. Comparisons are made to experimental data for Mach 2.5, 3.5, and 4.5. The accuracy of predicted aerodynamic coefficients is examined. Overpredictions in axial force and drag are explained by a closer examination of the surface pressure. These findings document sensitivities and best practices for future RANS database development of Mars entry vehicles.

RANS↗

NASA NTRSDate not supplied

Exploring the Accuracy of RANS Simulations for Mars Entry Vehicles

Accurate yet inexpensive predictions of aerodynamic coefficients for Mars entry vehicles have remained a consistent challenge over the past five decades. Below Mach 6, drag on the backshell becomes significant and must be accurately predicted. Steady Reynolds-averaged Navier-Stokes (RANS) models are commonly used, despite their poor predictions of backshell pressure. While scale-resolving simulations have shown promise in the past decade, there is still a need for cheap, accurate RANS predictions for large aerodynamic databases. The Mars Science Laboratory (MSL) is used as a case study to examine predictive accuracy and known shortcomings for RANS predictions of Mars entry vehicles. Several different grid generation techniques are compared, including a comparison between prismatic boundary layer grids and fully unstructured, tetrahedral grids. Comparisons are made to experimental data for Mach 2.5, 3.5, and 4.5. The accuracy of predicted aerodynamic coefficients is examined. Overpredictions in axial force and drag are explained by a closer examination of the surface pressure. These findings document sensitivities and best practices for future RANS database development of Mars entry vehicles.

CFD↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗

NASA NTRS2024

Completion of Dragonfly PPF Testing in the NFAC 80x120-Foot Wind Tunnel

The Dragonfly entry, descent, and landing team (JHU-APL, NASA Langley Research Center, NASA Ames Research Center, Sikorsky Aircraft) recently completed testing in the National Full-Scale Aerodynamics Complex (NFAC) 80x120-Foot Wind Tunnel. The facility is located at NASA Ames Research Center and is operated by the U.S. Air Force's Arnold Engineering Development Complex. The test was designed to simulate conditions of Preparation for Powered (PPF), an approximately 10-minute period during which the Lander is posed in front of the Backshell at low subsonic airspeeds, all under the main parachute, and the rotors are used to null residual Lander yaw rates prior to release i. e. de-spin. The Lander (approximately 50% scale) that previously was tested in the NASA Langley 14x22 Wind Tunnel in 2023 was used for the NFAC test with a new Backshell designed and fabricated specifically for NFAC. This memorandum summarizes execution of the test: objective, facility and models, instrumentation and data products, test procedure, and completed test matrix. Other future documents will include a test report (JHU-APL) and documentation of data processing/analysis and computational fluid dynamics (CFD) comparisons to test data (Sikorsky).

Dragonfly↗

NASA NTRSDate not supplied

Reconstruction of Thermal Protection System Aeroheating using a Green’s Function Approach

Inverse heat transfer (IHT) techniques are often used to reconstruct the surface heating conditions on spacecraft thermal protection systems (TPS) during atmospheric entry. Current IHT techniques for entry spacecraft applications, however, demand substantial computational resources, and are impractical for analyses such as uncertainty quantification and real-time health monitoring. In this paper, a Green’s function sensor fusion approach is used to reconstruct the TPS surface aeroheating conditions on experimental spaceflight and ground test systems from collocated temperature and heat flux sensors embedded in the TPS. The algorithm leverages Green’s functions to model the heat conduction within the spacecraft TPS and stabilizes the recovery of the surface heating condition using the direct heat flux sensor measurement. The algorithm is validated using arc-jet ground test data and applied to the reconstruction of the Mars 2020 backshell heating during Martian atmospheric entry. The performance of the algorithm is benchmarked against a current state-of-the-art IHT framework, FIAT_Opt. The Green’s function-based reconstruction algorithm recovers the net hot-wall heat flux absorbed by the TPS and the incident heat flux from the atmospheric entry environment in close agreement with FIAT_Opt. Notably, computation of the surface heating condition is completed in three orders of magnitude less time with the Green’s function sensor fusion approach using a consumer-grade PC, versus with FIAT_Opt running on a high performance computer cluster. The efficiency of the algorithm is leveraged to compute the uncertainty contributions of input parameters to the total uncertainty in reconstructed Mars 2020 backshell heating for the full atmospheric entry heat pulse. The sensitivity analysis uncovers that, at different times throughout the entry heat pulse, uncertainties in the TPS specific heat, thermal conductivity, and emissivity are all dominant drivers of the reconstruction uncertainty. These results demonstrate Green’s functions and sensor-fusion techniques as promising IHT approaches to reconstruct atmospheric entry environments from TPS-embedded measurements, and highlight how these techniques may give access to post-flight analyses previously hindered by the prohibitive cost of current methods.

Kenneth McAfee↗

NASA NTRSDate not supplied

Reconstruction of Thermal Protection System Aeroheating using a Green’s Function Approach

Inverse heat transfer (IHT) techniques are often used to reconstruct the surface heating conditions on spacecraft thermal protection systems (TPS) during atmospheric entry. Current IHT techniques for entry spacecraft applications, however, demand substantial computational resources, and are impractical for analyses such as uncertainty quantification and real-time health monitoring. In this paper, a Green’s function sensor fusion approach is used to reconstruct the TPS surface aeroheating conditions on experimental spaceflight and ground test systems from collocated temperature and heat flux sensors embedded in the TPS. The algorithm leverages Green’s functions to model the heat conduction within the spacecraft TPS and stabilizes the recovery of the surface heating condition using the direct heat flux sensor measurement. The algorithm is validated using arc-jet ground test data and applied to the reconstruction of the Mars 2020 backshell heating during Martian atmospheric entry. The performance of the algorithm is benchmarked against a current state-of-the-art IHT framework, FIAT_Opt. The Green’s function-based reconstruction algorithm recovers the net hot-wall heat flux absorbed by the TPS and the incident heat flux from the atmospheric entry environment in close agreement with FIAT_Opt. Notably, computation of the surface heating condition is completed in three orders of magnitude less time with the Green’s function sensor fusion approach using a consumer-grade PC, versus with FIAT_Opt running on a high performance computer cluster. The efficiency of the algorithm is leveraged to compute the uncertainty contributions of input parameters to the total uncertainty in reconstructed Mars 2020 backshell heating for the full atmospheric entry heat pulse. The sensitivity analysis uncovers that, at different times throughout the entry heat pulse, uncertainties in the TPS specific heat, thermal conductivity, and emissivity are all dominant drivers of the reconstruction uncertainty. These results demonstrate Green’s functions and sensor-fusion techniques as promising IHT approaches to reconstruct atmospheric entry environments from TPS-embedded measurements, and highlight how these techniques may give access to post-flight analyses previously hindered by the prohibitive cost of current methods.

Kenneth McAfee↗

NASA NTRS1995

Data Base On Cables And Connectors

Report describes Connector Adapter Cable Information Data Base (CONNAID) computer program, managing data base containing necessary information concerning electrical connectors, breakout boxes, adapter cables, backshells, and pertinent torque specifications for engineering project.

Bowen, Arlen R.↗

NASA NTRS2003

Aerogel Composites for Aerospace Thermal Protection

Aerogel composites formed by infiltrating organic and/or inorganic aerogels into fiber matrix materials enable us to exploit the low thermal conductivity and low density of aerogels while maintaining the strength, structure and other useful properties of a porous fiber matrix. New materials for extreme heating ranges are needed to insulate future spacecraft against the extreme heat of planetary atmospheric entry, but the insulation mass must be minimized in order to maximize the payload. A reusable system passively insulates to survive heating unchanged for relatively low heating. Ablators, which sacrifice mass to control heating, are used to protect vehicles against more extreme heating for a single use thermal protection system (TPS). Aerogel composites were fabricated and tested for spacecraft thermal protection. The high-temperaturey high heat flux tests described in this paper were performed in NASA Ames arc-jet facilities to simulate spacecraft atmospheric entry, and include heating conditions predicted for the forebody and backshell of the Mars Science Lander (MSL) entry probe. The aerogel composites tested showed excellent thermal performance in the arc-jet tests, functioning both as reusuable insulation under lower heat fluxes, and as ablative aerogels under the extreme heating predicted for the MSL forebody.

White, Susan↗

NASA NTRS2004

Mars Exploration Rover Terminal Descent Mission Modeling and Simulation

Because of NASA's added reliance on simulation for successful interplanetary missions, the MER mission has developed a detailed EDL trajectory modeling and simulation. This paper summarizes how the MER EDL sequence of events are modeled, verification of the methods used, and the inputs. This simulation is built upon a multibody parachute trajectory simulation tool that has been developed in POST I1 that accurately simulates the trajectory of multiple vehicles in flight with interacting forces. In this model the parachute and the suspended bodies are treated as 6 Degree-of-Freedom (6 DOF) bodies. The terminal descent phase of the mission consists of several Entry, Descent, Landing (EDL) events, such as parachute deployment, heatshield separation, deployment of the lander from the backshell, deployment of the airbags, RAD firings, TIRS firings, etc. For an accurate, reliable simulation these events need to be modeled seamlessly and robustly so that the simulations will remain numerically stable during Monte-Carlo simulations. This paper also summarizes how the events have been modeled, the numerical issues, and modeling challenges.

Raiszadeh, Behzad↗

NASA NTRS2006

Planetary Airplane Extraction System Development and Subscale Testing

The Aerial Regional-scale Environmental Survey (ARES) project employs an airplane as the science platform from which to collect science data in the previously inaccessible, thin atmosphere of Mars. In order for the airplane to arrive safely in the Martian atmosphere a number of sequences must occur. A critical element in the entry sequence at Mars is an extraction maneuver to separate the airplane quickly (in less than a second) from its protective backshell to reduce the possibility of re-contact, potentially leading to mission failure. This paper describes the development, testing, and lessons learned from building a 1/3 scale model of this airplane extraction system. This design, based on the successful Mars Exploration Rover (MER) extraction mechanism, employs a series of trucks rolling along tracks located on the surface of the central parachute can. Numerous tests using high speed video were conducted at the Langley Research Center (LaRC) to validate this concept. One area of concern was that that although the airplane released cleanly, a pitching moment could be introduced. While targeted for a Mars mission, this concept will enable environmental surveys by aircraft in other planetary bodies with a sensible atmosphere such as Venus or Saturn s moon, Titan.

Teter, John E., Jr.↗

NASA NTRS2006

Planetary Airplane Extraction System Development and Subscale Testing

The Aerial Regional-scale Environmental Survey (ARES) project will employ an airplane as the science platform from which to collect science data in the previously inaccessible, thin atmosphere of Mars. In order for the airplane to arrive safely in the Martian atmosphere, a number of sequences must occur. A critical element in the entry sequence at Mars is an extraction maneuver to separate the airplane quickly (in less than a second) from its protective backshell to reduce the possibility of re-contact, potentially leading to mission failure. This paper describes the development, testing, and lessons learned from building a 1/3 scale model of this airplane extraction system. This design, based on the successful Mars Exploration Rover (MER) extraction mechanism, employs a series of trucks rolling along tracks located on the surface of the central parachute can. Numerous tests using high speed video were conducted at the Langley Research Center to validate this concept. One area of concern was that that although the airplane released cleanly, a pitching moment could be introduced. While targeted for a Mars mission, this concept will enable environmental surveys by aircraft in other planetary bodies with a sensible atmosphere such as Venus or Saturn's moon, Titan.

Teter, John E., Jr.↗

NASA NTRS2006

Apollo Seals: A Basis for the Crew Exploration Vehicle Seals

The National Aeronautics and Space Administration is currently designing the Crew Exploration Vehicle (CEV) as a replacement for the Space Shuttle for manned missions to the International Space Station, as a command module for returning astronauts to the moon, and as an earth reentry vehicle for the final leg of manned missions to the moon and Mars. The CEV resembles a scaled-up version of the heritage Apollo vehicle; however, the CEV seal requirements are different than those from Apollo because of its different mission requirements. A review is presented of some of the seals used on the Apollo spacecraft for the gap between the heat shield and backshell and for penetrations through the heat shield, docking hatches, windows, and the capsule pressure hull.

Finkbeiner, Joshua R.↗