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  • Spacecraft Design, Testing and Performance  (21)
  • Size-at-age  (19)
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  • 1
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    Mars Microprobe Entry Analysis (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The Mars Microprobe mission will provide the first opportunity for subsurface measurements, including water detection, near the south pole of Mars. In this paper, performance of the Microprobe aeroshell design is evaluated through development of a six-degree-of-freedom (6-DOF) aerodynamic database and flight dynamics simulation. Numerous mission uncertainties are quantified and a Monte-Carlo analysis is performed to statistically assess mission performance. Results from this 6-DOF Monte-Carlo simulation demonstrate that, in a majority of the cases (approximately 2-sigma), the penetrator impact conditions are within current design tolerances. Several trajectories are identified in which the current set of impact requirements are not satisfied. From these cases, critical design parameters are highlighted and additional system requirements are suggested. In particular, a relatively large angle-of-attack range near peak heating is identified.
    Keywords: Spacecraft Design, Testing and Performance
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Integrated Design System (IDS) Tools for the Spacecraft Aeroassist/Entry Vehicle Design Process (1997)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: The definition of the Integrated Design System technology focus area as presented in the NASA Information Technology center of excellence strategic plan is described. The need for IDS tools in the aeroassist/entry vehicle design process is illustrated. Initial and future plans for spacecraft IDS tool development are discussed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Aeroassist Technology Workshop; Jan 14, 1997 - Jan 16, 1997; Pasadena, CA; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Mars Ascent Vehicle Flight Analysis (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The scientific objective of the Mars Surveyor Program 2005 mission is to return Mars rock, soil, and atmospheric samples to Earth for detailed analysis. The present investigation focuses on design of Mars Ascent Vehicle for this mission. Aerodynamic, aerothermodynamic, and trajectory design considerations are addressed to assess the ascent configuration, determine aerodynamic stability, characterize thermal protection system requirements, and ascertain the required system mass. Aerodynamic analysis reveals a subsonic static instability with the baseline configuration; however, stability augmentation options are proposed to mitigate this problem. The ascent aerothermodynamic environment is shown to be benign (on the order of the sea-level boiling point of water on Earth). As a result of these low thermal and pressure loads, a lightweight, low rigidity material can be employed as the aftbody aerodynamic shroud. The required nominal MAV lift-off mass is 426 kg for a December 2006 equatorial launch into a 300-km circular orbit with 30-degree inclination. Off-nominal aerodynamic and atmospheric conditions are shown to increase this liftoff mass by approximately 10%. Through performance of these analyses, the Mars Ascent Vehicle is deemed feasible with respect to the current mission mass and size constraints.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 98-2850 , 7th AIAA/ASME Joint Thermophysics and Heat Transfer Conference; Jun 15, 1998 - Jun 18, 1998; Albuquerque, NM; United States
    Format: application/pdf
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  • 4
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    Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Postflight analysis of the Mars Pathfinder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of preflight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement of the sonic line from the shoulder to the nose and back was confirmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coefficient. The sensitivity of the moment coefficient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with flight data. The atmospheric density as derived from axial coefficient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper98-2445 , 7th AIAA/ASME Joint Thermophysics and Heat Transfer Conference; Jun 15, 1998 - Jun 18, 1998; Albuquerque, NM; United States
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  • 5
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    Six Degree-of-Freedom Entry Dispersion Analysis for the METEOR Recovery Module (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The present study performs a six degree-of-freedom entry dispersion analysis for the Multiple Experiment Transporter to Earth Orbit and Return (METEOR) mission. METEOR offered the capability of flying a recoverable science package in a microgravity environment. However, since the Recovery Module has no active control system, an accurate determination of the splashdown position is difficult because no opportunity exists to remove any errors. Hence, uncertainties in the initial conditions prior to deorbit burn initiation, during deorbit burn and exo-atmospheric coast phases, and during atmospheric flight impact the splashdown location. This investigation was undertaken to quantify the impact of the various exo-atmospheric and atmospheric uncertainties. Additionally, a Monte-Carlo analysis was performed to statistically assess the splashdown dispersion footprint caused by the multiple mission uncertainties. The Monte-Carlo analysis showed that a 3-sigma splashdown dispersion footprint with axes of 43.3 nm (long), -33.5 nm (short), and 10.0 nm (crossrange) can be constructed. A 58% probability exists that the Recovery Module will overshoot the nominal splashdown site.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA-TM-112913 , NAS 1.26:112913 , AIAA Paper 96-0903 , Aerospace Sciences; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
    Format: application/pdf
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  • 6
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    A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration (2016)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA SciTech Conference; Jan 04, 2016 - Jan 08, 2016; San Diego, CA; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration (2016)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx.75 t that could deliver approx.28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx.24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx.20 t of propellant (including residuals) and the descent stage would require approx.21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA SciTech Conference; Jan 04, 2016 - Jan 08, 2016; San Diego, CA; United States
    Format: text
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  • 8
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    Two-Dimensional Finite Element Ablative Thermal Response Analysis of an Arcjet Stagnation Test (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The finite element ablation and thermal response (FEAtR, hence forth called FEAR) design and analysis program simulates the one, two, or three-dimensional ablation, internal heat conduction, thermal decomposition, and pyrolysis gas flow of thermal protection system materials. As part of a code validation study, two-dimensional axisymmetric results from FEAR are compared to thermal response data obtained from an arc-jet stagnation test in this paper. The results from FEAR are also compared to the two-dimensional axisymmetric computations from the two-dimensional implicit thermal response and ablation program under the same arcjet conditions. The ablating material being used in this arcjet test is phenolic impregnated carbon ablator with an LI-2200 insulator as backup material. The test is performed at the NASA, Ames Research Center Interaction Heating Facility. Spatially distributed computational fluid dynamics solutions for the flow field around the test article are used for the surface boundary conditions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2011-3617 , NF1676L-11639 , 42nd AIAA Thermophysics Conference; Jun 27, 2011 - Jun 30, 2011; Honolulu, HI; United States
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  • 9
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    A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration (2016)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: No abstract available
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA SciTech Conference; Jan 04, 2016 - Jan 08, 2016; San Diego, CA; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 10
    facet.materialart.
    Unknown
    A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration (2016)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx. 75 t that could deliver approx. 28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx. 24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx. 20 t of propellant (including residuals) and the descent stage would require approx. 21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA SciTech 2016; Jan 04, 2016 - Jan 06, 2016; San Diego, CA; United States
    Format: text
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