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  • 1
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    Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept (2002)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjet/scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.
    Keywords: Launch Vehicles and Launch Operations
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 321-335; CPIA-Publ-713-Vol-1
    Format: text
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  • 2
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    Unknown
    Status of the RBCC Direct-Connect Mixer Combustor Experiment (2002)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: The NASA Glenn Research Center is developing hydrogen based combined cycle propulsion technology for a single-stage-to-orbit launch vehicle application under a project called GTX. Rocket Based Combined Cycle (RBCC) propulsion systems incorporate one or more rocket engines into an airbreathing flow path to increase specific impulse as compared to an all rocket-powered vehicle. In support of this effort, an RBCC direct-connect test capability was established at the Engine Components Research Laboratory to investigate low speed, ejector ramjet, and initial ramjet operations and performance. The facility and test article enables the evaluation of two candidate low speed operating schemes; the simultaneous mixing and combustion (SMC) and independent ramjet stream (IRS). The SMC operating scheme is based on the fuel rich operations of the rocket where performance depends upon mixing between the rocket plume and airstream. In contrast, the IRS scheme fuels the airstream separately and uses the rocket plume to ignite the fuel-air mixture. This paper describes the test hardware and facility upgrades installed to support the RBCC tests. It also defines and discusses low speed technical challenges being addressed by the experiments. Finally, preliminary test results, including rocket risk mitigating tests, unfueled airflow tests, and the integrated system hot fire test will be presented.
    Keywords: Launch Vehicles and Launch Operations
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 125-134; CPIA-Publ-713-Vol-1
    Format: text
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  • 3
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    Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle (2008)
    In:  CASI
    Details
    Publication Date: 2018-06-12
    Description: As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst-case scenario occurs just before Upper Stage Main Engine Cut-Off when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.
    Keywords: Mechanical Engineering
    Type: 39th Aerospace Mechanisms Symposium; 345-358; NASA/CP-2008-215252
    Format: application/pdf
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  • 4
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    Buckling Testing and Analysis of Honeycomb Sandwich Panel Arc Segments of a Full-Scale Fairing Barrel: Comparison of In- and Out-of-Autoclave Facesheet Configurations (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Four honeycomb sandwich panels, representing 1/16th arc segments of a 10-m diameter barrel section of the Heavy Lift Launch Vehicle, were manufactured and tested under the NASA Composites for Exploration and the NASA Constellation Ares V programs. Two configurations were chosen for the panels: 6-ply facesheets with 1.125 in. honeycomb core and 8-ply facesheets with 1.0 in. honeycomb core. Additionally, two separate carbon fiber/epoxy material systems were chosen for the facesheets: in-autoclave IM7/977-3 and out-of-autoclave T40-800b/5320-1. Smaller 3 ft. by 5 ft. panels were cut from the 1/16th barrel sections and tested under compressive loading. Furthermore, linear eigenvalue and geometrically nonlinear finite element analyses were performed to predict the compressive response of each 3 ft. by 5 ft. panel. To improve the robustness of the geometrically nonlinear finite element model, measured surface imperfections were included in the geometry of the model. Both the linear and nonlinear models yielded good qualitative and quantitative predictions. Additionally, it was correctly predicted that the panel would fail in buckling prior to failing in strength. Furthermore, several imperfection studies were performed to investigate the influence of geometric imperfections, fiber angle misalignments, and three-dimensional effects on the compressive response of the panel.
    Keywords: Structural Mechanics
    Type: GRC-E-DAA-TN12569 , AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Jan 13, 2014 - Jan 19, 2014; National Harbor, MD; United States
    Format: application/pdf
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  • 5
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    Composite Payload Fairing Structural Architecture Assessment and Selection (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper provides a summary of the structural architecture assessments conducted and a recommendation for an affordable high performance composite structural concept to use on the next generation heavy-lift launch vehicle, the Space Launch System (SLS). The Structural Concepts Element of the Advanced Composites Technology (ACT) project and its follow on the Lightweight Spacecraft Structures and Materials (LSSM) project was tasked with evaluating a number of composite construction technologies for specific Ares V components: the Payload Shroud, the Interstage, and the Core Stage Intertank. Team studies strived to address the structural challenges, risks and needs for each of these vehicle components. Leveraging off of this work, the subsequent Composites for Exploration (CoEx) effort is focused on providing a composite structural concept to support the Payload Fairing for SLS. This paper documents the evaluation and down selection of composite construction technologies and evolution to the SLS Payload Fairing. Development of the evaluation criteria (also referred to as Figures of Merit or FOMs), their relative importance, and association to vehicle requirements are presented. A summary of the evaluation results, and a recommendation of the composite concept to baseline in the Composites for Exploration (CoEx) project is presented. The recommendation for the SLS Fairing is a Honeycomb Sandwich architecture based primarily on affordability and performance with two promising alternatives, Hat stiffened and Fiber Reinforced Foam (FRF) identified for eventual program block upgrade.
    Keywords: Spacecraft Design, Testing and Performance
    Type: E-18094 , Society for the Advancement of Material and Process Engineering (SAMPE) conference; May 21, 2012 - May 24, 2012; Baltimore, MD; United States
    Format: application/pdf
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  • 6
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    Unknown
    Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept (2003)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic air- breathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demon- strate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA/TM-2003-212315 , E-13316 , NAS 1.15:212315 , APS-III-22 , Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittees Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
    Format: application/pdf
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  • 7
    facet.materialart.
    Unknown
    Status of the RBCC Direct-Connect Mixer Combustor Experiment (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The NASA Glenn Research Center is developing hydrogen based combined cycle propulsion technology for a single-stage-to-orbit launch vehicle application under a project called GTX. Rocket Based Combined Cycle (RBCC) propulsion systems incorporate one or more rocket engines into an airbreathing flow path to increase specific impulse as compared to an all rocket-powered vehicle. In support of this effort, an RBCC direct-connect test capability was established at the Engine Components Research Laboratory to investigate low speed, ejector ramjet, and initial ramjet operations and performance. The facility and test article enables the evaluation of two candidate low speed operating schemes; the simultaneous mixing and combustion (SMC) and independent ramjet stream (IRS). The SMC operating scheme is based on the fuel rich operations of the rocket where performance depends upon mixing between the rocket plume and airstream. In contrast, the IRS scheme fuels the airstream separately and uses the rocket plume to ignite the fuel-air mixture. This paper describes the test hardware and facility upgrades installed to support the RBCC tests. It also defines and discusses low speed technical challenges being addressed by the experiments. Finally, preliminary test results, including rocket risk mitigating tests, unfueled airflow tests, and the integrated system hot fire test will be presented.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211555 , NAS 1.15:211555 , E-13334 , Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation Subcommittes Joint Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
    Format: application/pdf
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  • 8
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    Unknown
    Lessons Learned from the Construction of Upgrades to the NASA Glenn Icing Research Tunnel and Re-activation Testing (2001)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Major upgrades were made in 1999 to the 6- by 9-Foot (1.8- by 2.7-m) Icing Research Tunnel (IRT) at the NASA Glenn Research Center. These included replacement of the electronic controls for the variable-speed drive motor, replacement of the heat exchanger, complete replacement and enlargement of the leg of the tunnel containing the new heat-exchanger, the addition of flow-expanding and flow-contracting turning vanes upstream and downstream of the heat exchanger, respectively, and the addition of fan outlet guide vanes (OGV's). This paper presents an overview of the construction and reactivation testing phases of the project. Important lessons learned during the technical and contract management work are documented.
    Keywords: Research and Support Facilities (Air)
    Type: NASA/TM-2001-210688 , NAS 1.15:210688 , E-12615 , AIAA Paper 2001-0231 , 39th Aerospace Sciences Meeting and Exhibit; Jan 08, 2001 - Jan 11, 2001; Reno, NV; United States
    Format: application/pdf
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  • 9
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    Unknown
    Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and it's engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.
    Keywords: Launch Vehicles and Launch Operations
    Type: 39th Aerospace Mechanisms Symposium; May 07, 2008 - May 09, 2008; Huntsville, AL; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
    facet.materialart.
    Unknown
    Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2008-215182 , E-16305-1 , 35th Aerospace Mechanisms Symposium; May 07, 2008 - May 09, 2008; Huntsville, AL; United States
    Format: application/pdf
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