SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Pressure data are presented in tabular form on the external tank base and SRM booster base of the space shuttle launch vehicle configuration model. For Vol. 3, see N75-23657.
Aerodynamic characteristics determined during development of Apollo launch escape vehicle configuration in wind tunnel tests
Linear aerodynamic load distributions on missile and launch vehicle configurations
Wind tunnel investigation of aerodynamic pressures on Apollo launch escape vehicle configuration
Tip tank configuration concept for integrated launch reentry vehicles
Mechanical design integration of Saturn S- IVB STAGE of Saturn IB and Saturn V launch vehicle configuration
Effects of simulated jet exhaust flow on transonic aerodynamic characteristics of Apollo launch escape vehicle models
Results of tests conducted on an 0.030-scale launch configuration model of the space shuttle vehicle 140A/B in the NASA/ARC 11-foot unitary plan wind tunnel are presented. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 1.4. Surface pressure distributions were obtained simultaneously with six-component stability and control force data on the complete launch configuration. The configuration consisted of the orbiter, an external tank, two solid rocket boosters, and associated intercomponent attach hardware. Angles of attack and sideslip from -10 degrees to +10 degrees were investigated.
The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a high-level analysis of various air launch vehicle configurations, objectively determining maximum launch vehicle payload while considering carrier aircraft capabilities and given dimensional constraints. With the renewed interest in aerial launch of low-earth orbit payloads, referenced by programs such as Stratolaunch and Spaceship2, there exists a need to qualify the boundaries of the trade space, identify performance envelopes, and understand advantages and limiting factors of designing for maximum payload capability. Using the NASA/DARPA Horizontal Launch Study (HLS) Point Design 2 (PD-2) as a pointof- departure configuration, two independent design actions were undertaken. Both designs utilized a Boeing 747-400F as the carrier aircraft, LOX/RP-1 first stage and LOX/LH2 second stage. Each design was sized to meet dimensional and mass constraints while optimizing propellant loads and stage delta V splits. All concepts, when fully loaded, exceeded the allowable Gross Takeoff Weight (GTOW) of the aircraft platform. This excess mass was evaluated as propellant/fuel offload available for a potential in-flight propellant loading scenario. Results indicate many advantages such as payload delivery of approximately 47,000 lbm and significant mission flexibility including variable launch site inclination and launch window. However, in-flight cryogenic fluid transfer and carrier aircraft platform integration are substantial technical hurdles to the realization of such a system configuration.
The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center conducted a high-level analysis of various air launch vehicle configurations, objectively determining maximum launch vehicle payload while considering carrier aircraft capabilities and given dimensional constraints. With the renewed interest in aerial launch of low-earth orbit payloads, referenced by programs such as Stratolaunch and Spaceship2, there existed a need to qualify the boundaries of the trade space, identify performance envelopes, and understand advantages and limiting factors of designing for maximum payload capability. Using the NASA/DARPA Horizontal Launch Study (HLS) Point Design 2 (PD-2) as a point-of-departure configuration, two independent design actions were undertaken. Both configurations utilized a Boeing 747-400F as the carrier aircraft, LOX/RP-1 first stage and LOX/LH2 second stage. Each design was sized to meet dimensional and mass constraints while optimizing propellant loads and stage delta V (ΔV) splits. All concepts, when fully loaded, exceeded the allowable Gross Takeoff Weight (GTOW) of the aircraft platform. This excess mass was evaluated as propellant/fuel offload available for a potential in-flight refueling scenario. Results indicate many advantages such as large, relative payload delivery of approximately 47,000 lbm and significant mission flexibility, such as variable launch site inclination and launch window; however, in-flight cryogenic fluid transfer and carrier aircraft platform integration are substantial technical hurdles to the realization of such a system configuration.
The purpose of the Advanced Transportation System Study (ATSS) task area 1 study effort is to examine manned launch vehicle booster concepts and two-way cargo transfer and return vehicle concepts to determine which of the many proposed concepts best meets NASA's needs for two-way transportation to low earth orbit. The study identified specific configurations of the normally unmanned, expendable launch vehicles (such as the National Launch System family) necessary to fly manned payloads. These launch vehicle configurations were then analyzed to determine the integrated booster/spacecraft performance, operations, reliability, and cost characteristics for the payload delivery and return mission. Design impacts to the expendable launch vehicles which would be required to perform the manned payload delivery mission were also identified. These impacts included the implications of applying NASA's man-rating requirements, as well as any mission or payload unique impacts. The booster concepts evaluated included the National Launch System (NLS) family of expendable vehicles and several variations of the NLS reference configurations to deliver larger manned payload concepts (such as the crew logistics vehicle (CLV) proposed by NASA JSC). Advanced, clean sheet concepts such as an F-1A engine derived liquid rocket booster (LRB), the single stage to orbit rocket, and a NASP-derived aerospace plane were also included in the study effort. Existing expendable launch vehicles such as the Titan 4, Ariane 5, Energia, and Proton were also examined. Although several manned payload concepts were considered in the analyses, the reference manned payload was the NASA Langley Research Center's HL-20 version of the personnel launch system (PLS). A scaled up version of the PLS for combined crew/cargo delivery capability, the HL-42 configuration, was also included in the analyses of cargo transfer and return vehicle (CTRV) booster concepts. In addition to strictly manned payloads, two-way cargo transportation systems (CTRV's) were also examined. The study provided detailed design and analysis of the performance, reliability, and operations of these concepts. The study analyzed these concepts as unique systems and also analyzed several combined CTRV/booster configurations as integrated launch systems (such as for launch abort analyses). Included in the set of CTRV concepts analyzed were the medium CTRV, the integral CTRV (in both a pressurized and unpressurized configuration), the winged CTRV, and an attached cargo carrier for the PLS system known as the PLS caboose.
Ground wind oscillatory loads and launch vehicle configurations
The objective of this study is to validate a computational methodology for the aerodynamic performance of an advanced conical launch vehicle configuration. The computational methodology is based on a three-dimensional, viscous flow, pressure-based computational fluid dynamics formulation. Both wind-tunnel and ascent flight-test data are used for validation. Emphasis is placed on multiple-engine power-on effects. Computational characterization of the base drag in the critical subsonic regime is the focus of the validation effort; until recently, almost no multiple-engine data existed for a conical launch vehicle configuration. Parametric studies using high-order difference schemes are performed for the cold-flow tests, whereas grid studies are conducted for the flight tests. The computed vehicle axial force coefficients, forebody, aftbody, and base surface pressures compare favorably with those of tests. The results demonstrate that with adequate grid density and proper distribution, a high-order difference scheme, finite rate afterburning kinetics to model the plume chemistry, and a suitable turbulence model to describe separated flows, plume/air mixing, and boundary layers, computational fluid dynamics is a tool that can be used to predict the low-speed aerodynamic performance for rocket design and operations.
Performance characteristics of Aerobee sounding rocket and launch vehicle configuration
The feasibility of separating a two-stage winged, parallel configured launch vehicle at Mach 3 has been determined. The two-stage fully-reusable launch vehicle is comprised of an orbiter, which can be used alone as a single-stage-to-orbit vehicle, and a booster which is identical to the baseline orbiter design. A three-degree-of-freedom analysis was conducted on the ascent trajectory of the two-stage concept to determine the optimum staging conditions. The booster provides all propellant for both stages using a crossfeed system until staging occurs. After staging, the empty booster glides back to the launch site for a horizontal landing. To analyze the separation maneuver, the six-degree-of-freedom equations of motion for both stages were numerically integrated using a computer program which incorporates analytically derived interference aerodynamic data. The necessary conditions for a successful separation were determined and required the use of active controls on both the orbiter and booster. Once the booster and orbiter were separated, a three-degree-of-freedom analysis was conducted on the booster's glideback trajectory.
Transonic wind tunnel study of ground wind loads on Saturn V and other launch vehicle configurations
Experimental aerodynamic investigations were conducted on a 6-OTS 0.015-scale model. The Ames dual sting support separation rig was used to obtain grid-type data for tank-booster abort from orbiter (SSV). Freestream data were obtained for the orbiter to provide a baseline for evaluation of proximity effects. Data were obtained at Mach numbers from 0.32 to 1.1, and Reynolds number per foot varying from 2.1 million to 3.9 million. Data are not presented. Because of balance failure, a very substantial portion of the test was run with a dummy balance in the tank boosters configuration.
The results of the tests are presented along with aerodynamic loads data obtained at Mach numbers from 0.6 to 1.4. Surface pressure distributions were obtained simultaneously with six-component stability and control force data on the complete launch configuration. The configuration consisted of the orbiter, an external tank, two solid rocket boosters, and associated intercomponent attach hardware. Angles of attack and sideslip from -10 deg to +10 deg were investigated. The tests were conducted from 4 September 1973 through 13 September 1973.