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  • 1
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    Thermal Analysis on Plume Heating of the Main Engine on the Crew Exploration Vehicle Service Module (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: The crew exploration vehicle (CEV) service module (SM) main engine plume heating is analyzed using multiple numerical tools. The chemical equilibrium compositions and applications (CEA) code is used to compute the flow field inside the engine nozzle. The plume expansion into ambient atmosphere is simulated using an axisymmetric space-time conservation element and solution element (CE/SE) Euler code, a computational fluid dynamics (CFD) software. The thermal analysis including both convection and radiation heat transfers from the hot gas inside the engine nozzle and gas radiation from the plume is performed using Thermal Desktop. Three SM configurations, Lockheed Martin (LM) designed 604, 605, and 606 configurations, are considered. Design of multilayer insulation (MLI) for the stowed solar arrays, which is subject to plume heating from the main engine, among the passive thermal control system (PTCS), are proposed and validated.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2007-215049 , TFAWS 07-1012 , E-16260 , Thermal and Fluids Analysis Workshop (TFAWS) 2007; 10-14 Seo, 2007; Warrensville Heights, OH; United States
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  • 2
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    Development of Parametric Mass and Volume Models for an Aerospace SOFC/Gas Turbine Hybrid System (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In aerospace power systems, mass and volume are key considerations to produce a viable design. The utilization of fuel cells is being studied for a commercial aircraft electrical power unit. Based on preliminary analyses, a SOFC/gas turbine system may be a potential solution. This paper describes the parametric mass and volume models that are used to assess an aerospace hybrid system design. The design tool utilizes input from the thermodynamic system model and produces component sizing, performance, and mass estimates. The software is designed such that the thermodynamic model is linked to the mass and volume model to provide immediate feedback during the design process. It allows for automating an optimization process that accounts for mass and volume in its figure of merit. Each component in the system is modeled with a combination of theoretical and empirical approaches. A description of the assumptions and design analyses is presented.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2005-213819 , GT2005-68334 , E-15177 , Turbo Expo 2005; Jun 06, 2005 - Jun 09, 2005; Reno, NV; United States
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  • 3
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    Orion Service Module Reaction Control System Plume Impingement Analysis Using PLIMP/RAMP2 (2009)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Orion Crew Exploration Vehicle Service Module Reaction Control System engine plume impingement was computed using the plume impingement program (PLIMP). PLIMP uses the plume solution from RAMP2, which is the refined version of the reacting and multiphase program (RAMP) code. The heating rate and pressure (force and moment) on surfaces or components of the Service Module were computed. The RAMP2 solution of the flow field inside the engine and the plume was compared with those computed using GASP, a computational fluid dynamics code, showing reasonable agreement. The computed heating rate and pressure using PLIMP were compared with the Reaction Control System plume model (RPM) solution and the plume impingement dynamics (PIDYN) solution. RPM uses the GASP-based plume solution, whereas PIDYN uses the SCARF plume solution. Three sets of the heating rate and pressure solutions agree well. Further thermal analysis on the avionic ring of the Service Module showed that thermal protection is necessary because of significant heating from the plume.
    Keywords: Space Transportation and Safety
    Type: E-16823 , 47th AIAA Aerospace Sciences Meeting and Exhibit; Jan 05, 2009 - Jan 08, 2009; Florida; United States
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  • 4
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    Thermal Analysis on Cryogenic Liquid Hydrogen Tank on an Unmanned Aerial Vehicle System (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Thermal analyses are performed on the liquid hydrogen (LH2) tank designed for an unmanned aerial vehicle (UAV) powered by solar arrays and a regenerative proton-exchange membrane (PEM) fuel cell. A 14-day cruise mission at a 65,000 ft altitude is considered. Thermal analysis provides the thermal loads on the tank system and the boiling-off rates of LH2. Different approaches are being considered to minimize the boiling-off rates of the LH2. It includes an evacuated multilayer insulation (MLI) versus aerogel insulation on the LH2 tank and aluminum versus stainless steel spacer rings between the inner and outer tank. The resulting boil-off rates of LH2 provided by the one-dimensional model and three-dimensional finite element analysis (FEA) on the tank system are presented and compared to validate the results of the three-dimensional FEA. It concludes that heat flux through penetrations by conduction is as significant as that through insulation around the tank. The tank system with MLI insulation and stainless steel spacer rings result in the lowest boiling-off rate of LH2.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2007-214675 , AIAA Paper 2007-1218 , E-15811 , 45th Aerospace Sciences Meeting and Exhibit; Jan 08, 2007 - Jan 11, 2007; Reno, NV; United States
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  • 5
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    CFD Simulation on the J-2X Engine Exhaust in the Center-Body Diffuser and Spray Chamber at the B-2 Facility (2009)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A computational fluid dynamic (CFD) code is used to simulate the J-2X engine exhaust in the center-body diffuser and spray chamber at the Spacecraft Propulsion Facility (B-2). The CFD code is named as the space-time conservation element and solution element (CESE) Euler solver and is very robust at shock capturing. The CESE results are compared with independent analysis results obtained by using the National Combustion Code (NCC) and show excellent agreement.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2009-215464 , TFAWS08-1013 , E-16610
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  • 6
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    Conjugate Heat Transfer Analyses on the Manifold for Ramjet Fuel Injectors (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Three-dimensional conjugate heat transfer analyses on the manifold located upstream of the ramjet fuel injector are performed using CFdesign, a finite-element computational fluid dynamics (CFD) software. The flow field of the hot fuel (JP-7) flowing through the manifold is simulated and the wall temperature of the manifold is computed. The three-dimensional numerical results of the fuel temperature are compared with those obtained using a one-dimensional analysis based on empirical equations, and they showed a good agreement. The numerical results revealed that it takes around 30 to 40 sec to reach the equilibrium where the fuel temperature has dropped about 3 F from the inlet to the exit of the manifold.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2006-214455 , E-15719 , TFAWS06-1036
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  • 7
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    Engineering Analysis Studies for Preliminary Design of Lightweight Cryogenic Hydrogen Tanks in UAV Applications (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A series of engineering analysis studies were conducted to investigate the potential application of nanoclay-enhanced graphite/epoxy composites and polymer cross-linked silica aerogels in cryogenic hydrogen storage tank designs. This assessment focused on the application of these materials in spherical tank designs for unmanned aeronautic vehicles with mission durations of 14 days. Two cryogenic hydrogen tank design concepts were considered: a vacuum-jacketed design and a sandwiched construction with an aerogel insulating core. Analyses included thermal and structural analyses of the tank designs as well as an analysis of hydrogen diffusion to specify the material permeability requirements. The analyses also provided material property targets for the continued development of cross-linked aerogels and nanoclay-enhanced graphite/epoxy composites for cryogenic storage tank applications. The results reveal that a sandwiched construction with an aerogel core is not a viable design solution for a 14-day mission. A vacuum-jacketed design approach was shown to be far superior to an aerogel. Aerogel insulation may be feasible for shorter duration missions. The results also reveal that the application of nanoclay-enhanced graphite/epoxy should be limited to the construction of outer tanks in a vacuum-jacketed design, since a graphite/epoxy inner tank does not provide a significant weight savings over aluminum and since the ability of nanoclay-enhanced graphite/epoxy to limit hydrogen permeation is still in question.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TP-2006-214094 , E-15431
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  • 8
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    Orion Service Module Reaction Control System Plume Impingement Analysis Using PLIMP/RAMP2 (2009)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Orion Crew Exploration Vehicle Service Module Reaction Control System engine plume impingement was computed using the plume impingement program (PLIMP). PLIMP uses the plume solution from RAMP2, which is the refined version of the reacting and multiphase program (RAMP) code. The heating rate and pressure (force and moment) on surfaces or components of the Service Module were computed. The RAMP2 solution of the flow field inside the engine and the plume was compared with those computed using GASP, a computational fluid dynamics code, showing reasonable agreement. The computed heating rate and pressure using PLIMP were compared with the Reaction Control System plume model (RPM) solution and the plume impingement dynamics (PIDYN) solution. RPM uses the GASP-based plume solution, whereas PIDYN uses the SCARF plume solution. Three sets of the heating rate and pressure solutions agree well. Further thermal analysis on the avionic ring of the Service Module was performed using MSC Patran/Pthermal. The obtained temperature results showed that thermal protection is necessary because of significant heating from the plume.
    Keywords: Avionics and Aircraft Instrumentation
    Type: NASA/TM-2009-215601 , AIAA-2009-0834 , E-16823-1 , 47th AIAA Aerospace Sciences Meeting and Exhibit; Jan 05, 2009 - Jan 08, 2009; Orlando, FL; United States
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  • 9
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    Hybrid Solid Oxide Fuel Cell/Gas Turbine System Design for High Altitude Long Endurance Aerospace Missions (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A system level analysis, inclusive of mass, is carried out for a cryogenic hydrogen fueled hybrid solid oxide fuel cell and bottoming gas turbine (SOFC/GT) power system. The system is designed to provide primary or secondary electrical power for an unmanned aerial vehicle (UAV) over a high altitude, long endurance mission. The net power level and altitude are parametrically varied to examine their effect on total system mass. Some of the more important technology parameters, including turbomachinery efficiencies and the SOFC area specific resistance, are also studied for their effect on total system mass. Finally, two different solid oxide cell designs are compared to show the importance of the individual solid oxide cell design on the overall system. We show that for long mission durations of 10 days or more, the fuel mass savings resulting from the high efficiency of a SOFC/GT system more than offset the larger powerplant mass resulting from the low specific power of the SOFC/GT system. These missions therefore favor high efficiency, low power density systems, characteristics typical of fuel cell systems in general.
    Keywords: Propellants and Fuels
    Type: NASA/TM-2006-214328 , E-15560 , FUELCELL2006-97095 , Fourth International ASME Conference on Fuel Cell Science, Engineering and Technology; Jun 19, 2006 - Jun 21, 2006; Irvine, CA; United States
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  • 10
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    Ascent Heating Thermal Analysis on the Spacecraft Adaptor (SA) Fairings and the Interface with the Crew Launch Vehicle (CLV) (2009)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: When the crew exploration vehicle (CEV) is launched, the spacecraft adaptor (SA) fairings that cover the CEV service module (SM) are exposed to aero heating. Thermal analysis is performed to compute the fairing temperatures and to investigate whether the temperatures are within the material limits for nominal ascent aero heating case. Heating rates from Thermal Environment (TE) 3 aero heating analysis computed by engineers at Marshall Space Flight Center (MSFC) are used in the thermal analysis. Both MSC Patran 2007r1b/Pthermal and C&R Thermal Desktop 5.1/Sinda models are built to validate each other. The numerical results are also compared with those reported by Lockheed Martin (LM) and show a reasonably good agreement.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2009-215474 , TFAWS08-1008 , E-16611 , 2008 Thermal and Fluids Analysis Workshop; Aug 18, 2008 - Aug 22, 2008; San Jose, CA; United States
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