SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 8 - April 21, 2006
18 SPACECRAFT DESIGN, TESTING AND PERFORMANCE
Includes satellites; space platforms; space stations; spacecraft systems and components such as thermal and environmental controls; and spacecraft control and stability characteristics.
For life support systems see 54 Man/System Technology and Life Support.
For related information see also 05 Aircraft Design, Testing and Performance; 39 Structural Mechanics; and 16 Space Transportation and Safety
20060010213 Mississippi Univ., MS, USA
Development of a New Approach to Instrument Model Design Used by Team X
Andrew, Shanna E.; Summer Student Research Presentations; August 2005, pp. 29; In English; See also 20060010186; No Copyright; Available from CASI only as part of the entire parent document
The Jet Propulsion Laboratory's Advanced Design Team was formed in April 1995 to improve the quality and reduce the cost of JPL proposals and advanced mission studies. Currently a consolidation attempt is underway to develop a Model Library for use by JPL's Advanced Projects Design Team by collecting existing instrument models for inclusion in the library. This will allow users to readily find models of interest. In addition to this, there is also an attempt underway to develop a new approach to instrument model design used by the Advanced Design Team (Team X). This new approach consists of splitting up the different model parts such as orbital parameters, instrument parameters and instrument outputs into separate searchable parts. The user can then decide between design trades and use the different pieces to construct a model that will fit their needs. As well, this will lead to the opportunity for the large variety of usable instrument models. Author
Cost Reduction; Jet Propulsion; Libraries; Consolidation
20060010218 North Dakota Univ., Grand Forks, ND, USA
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CoMET: Cost and Mass Evaluation Tool for Spacecraft and Mission Design
Bieber, Ben S.; Summer Student Research Presentations; August 2005, pp. 29; In English; See also 20060010186; No Copyright; Available from CASI only as part of the entire parent document
New technology in space exploration is often developed without a complete knowledge of its impact. While the immediate benefits of a new technology are obvious, it is harder to understand its indirect consequences, which ripple through the entire system. COMET is a technology evaluation tool designed to illuminate how specific technology choices affect a mission at each system level. COMET uses simplified models for mass, power, and cost to analyze performance parameters of technologies of interest. The sensitivity analysis that CoMET provides shows whether developing a certain technology will greatly benefit the project or not. CoMET is an ongoing project approaching a web-based implementation phase. This year, development focused on the models for planetary daughter craft, such as atmospheric probes, blimps and balloons, and landers. These models are developed through research into historical data, well established rules of thumb, and engineering judgment of experts at JPL. The model is validated by corroboration with JpL advanced mission studies. Other enhancements to COMET include adding launch vehicle analysis and integrating an updated cost model. When completed, COMET will allow technological development to be focused on areas that will most drastically improve spacecraft performance. Author
Atmospheric Sounding; Space Exploration; Cost Analysis; Environmental Monitoring; Spacecraft Design; Sensitivity Analysis; Mission Planning; Launch Vehicles
20060010235 Minnesota Univ., USA
Relevant Parameter Evolution in Team X Mission Concept Designs
Huelman, Brandon N.; Summer Student Research Presentations; August 2005, pp. 35-36; In English; See also 20060010186; No Copyright; Available from CASI only as part of the entire parent document
The Advanced Projects Design Team, also known as Team X, is a concurrent engineering team that quickly and cheaply designs space mission architectures including the flight system and subsystems, the trajectory, and ground system. Through the use of ICEMaker, an Excel spreadsheet database, the parameters from each subsystem can be shared and used among the other subsystems. This allows for entire missions to be planned with only a few short design team sessions. Based on the results, the feasibility of the mission concept can be determined. Over the Years since the team was created, the amount of information being shared among subsystems on the database has increased, however many of the parameters are now obsolete. Removal of these unused parameters Will clean UP the database and help to streamline the mission design process. By comparing parameter files from previous Team X mission studies, the parameter usage can be determined. As was initially suspected there are more unused parameters on the database than parameters that are actually used. Author
Mission Planning; Space Missions; Data Bases; Concurrent Engineering
20060010242 Wellesley Coll., MA, USA
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Team X Spacecraft Instrument Database Consolidation
Wallenstein, Kelly A.; Summer Student Research Presentations; August 2005, pp. 48; In English; See also 20060010186; No Copyright; Available from CASI only as part of the entire parent document
In the past decade, many changes have been made to Team X's process of designing each spacecraft, with the purpose of making the overall procedure more efficient over time. One such improvement is the use of information databases from previous missions, designs, and research. By referring to these databases, members of the design team can locate relevant instrument data and significantly reduce the total time they spend on each design. The files in these databases were stored in several different formats with various levels of accuracy. During the past 2 months, efforts have been made in an attempt to combine and organize these files. The main focus was in the Instruments department, where spacecraft subsystems are designed based on mission measurement requirements.Acommon database was developed for all instrument parameters using Microsoft Excel to minimize the time and confusion experienced when searching through files stored in several different formats and locations. By making this collection of information more organized, the files within them have become more easily searchable. Additionally, the new Excel database offers the option of importing its contents into a more efficient database management system in the future. This potential for expansion enables the database to grow and acquire more search features as needed. Author
Data Bases; Spacecraft Instruments; Spacecraft Design; Consolidation
20060010453 Boeing Co., Houston, TX, USA
Space Shuttle Partial Stack Rollout Test Analytical Correlation In Support Of Fatigue Load Development
DelBasso, Steve; Dolenz, Jim; Wilson, Lee; [2005]; 2 pp.; In English; International Modal Analysis Conference, 31 Jan. - 3 Feb. 2005, Orlando, FL, USA; No Copyright; Avail.: CASI: A01, Hardcopy
A rollout test with only the Solid Rocket Boosters was conducted in November 2003 to gather structural dynamic response data of the transportation environment from the Vehicle Assembly Building to the Launch Pad. The data was acquired to develop and validate analytical methods used to predict rollout Orbiter fatigue load spectra. Earlier predictions computed by a base drive approach with only 5 input drive degrees-of-freedom raised questions that commissioned the partial stack test. Not only was there a concern because of the input degree-of-freedom omission due to measurement limitations, but there was also a concern with the implementation of the 'large mass' itself. Three methods were evaluated with the partial stack test data. The analytical correlations to measured strain derived SRB base loads and acceleration5 showed the earlier 5 degree-offreedom base drive approach to yield the most conservative results for all quantities monitored except the SRB base moment about the axis in which the input drive was missing. This non-conservative shortcoming led to a recommendation to use either the 6 degree-of-freedom base drive or the 12 degree-of-freedom Craig-Bampton boundary drive methods whose results did not substantially differ. Author
Booster Rocket Engines; Degrees of Freedom; Dynamic Response; Load Tests; Space Shuttle Boosters; Solid Propellant Rocket Engines
20060010454 NASA Johnson Space Center, Houston, TX, USA
Rendezvous and Docking for Space Exploration
Machula, M. F.; Crain, T.; Sandhoo, G. S.; [2005]; 1 pp.; In English; 1st Space Exploration Conference, 30 Jan. - 1 Feb. 2005, Orlando, FL, USA; No Copyright; Avail.: Other Sources; Abstract Only
To achieve the exploration goals, new approaches to exploration are being envisioned that include robotic networks, modular systems, pre-positioned propellants and in-space assembly in Earth orbit, Lunar orbit and other locations around the cosmos. A fundamental requirement for rendezvous and docking to accomplish in-space assembly exists in each of these locations. While existing systems and technologies can accomplish rendezvous and docking in low earth orbit, and rendezvous and docking with crewed systems has been successfully accomplished in low lunar orbit, our capability must extend toward autonomous rendezvous and docking. To meet the needs of the exploration vision in-space assembly requiring both crewed and uncrewed vehicles will be an integral part of the exploration architecture. This paper focuses on the intelligent application of autonomous rendezvous and docking technologies to meet the needs of that architecture. It also describes key technology investments that will increase the exploration program's ability to ensure mission success, regardless of whether the rendezvous are fully automated or have humans in the loop. Author
Autonomous Docking; Spacecraft Docking; Space Exploration; Orbital Rendezvous; Low Earth Orbits; Lunar Orbits
20060010486 Research and Technology Organization, Neuilly-sur-Seine, France
Critical Technologies for Hypersonic Vehicle Development
December 2005; In English; AVT/VKI Lecture Series, 10-14 May 2004, Rhode-Saint-Genese, Belgium; See also 20060010487 - 20060010501 Report No.(s): RTO-EN-AVT-116; AC/323(AVT-116)TP/90; Copyright; Avail.: CASI: C01, CD-ROM
The objective of the present RTO/VKI Lecture Series is to focus attention on the technologies that play a critical role in hypersonic vehicle development. Owing to the demanding flight environment, successful development of hypersonic vehicles will require progress in understanding and surmounting specific technical problems in ground test, numerical simulation and flight test. Following the introductory remarks stating the objectives of the Lecture Series, a historical perspective on hypersonic vehicle development was given to establish the context and to highlight the emerging role of numerical simulation in vehicle design. A thorough review of ongoing vehicle development and flight-test programs completed the introductory portion. To provide sufficient background information, lectures were presented on compressible flows of perfect and imperfect gases, with special attention given to shock-wave boundary-layer interactions, laminar to turbulent transition, and high temperature gas properties. Advancements in computational modeling capabilities for hypersonic vehicle design were addressed, and the ongoing need for validation data, as well as a consistent validation methodology, where highlighted. Specific critical technology areas, including trajectory aerothermal environment definition, system trade studies, propulsion systems, aerodynamic control laws, flight instrumentation, thermal protection systems and flight experiments were all addressed in separate lectures given by international experts. Finally, the potential of plasma devices to ameliorate or avoid critical aerothermal phenomena was presented, providing a good example of the type of creative effort required for achieving progress. In all of the presentations the needs of future hypersonic flight vehicle development programs were addressed. Derived from text
Hypersonic Vehicles; Flight Simulation; Ground Tests; Flight Tests; High Temperature Gases; Shock Waves; Boundary Layers; Turbulence; Transition Temperature
20060010487 Johns Hopkins Univ., Laurel, MD, USA
Future Technologies - Application of Plasma Devices for Vehicle Systems
VanWie, D. M.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 15-1 - 15-40; In English; See also 20060010486; Original contains black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
The application of plasma devices for controlling and enhancing aerodynamic phenomena encountered in atmospheric flight vehicles is actively being investigated by many research groups throughout the world. These devices operate using the generation of plasma within the flowfield, generally through an electrical discharge or injection of an electron beam, and then using either the effects of the energy deposition itself or through application of electrostatic or Lorentz forces to modify the aerodynamic phenomena of interest. Plasma devices are being explored for applications ranging from drag reduction and steering control to boundary layer modification to ignition and combustion system enhancement. Note that since the plasma generation generally operates using electrically controlled phenomena, the frequency response of a control system can be very fast relative to hydraulic and electromechanical systems, so new control strategies can be explored. In the present notes, a variety of plasma flow control devices are described together with a discussion of their potential system-level impacts. High level results are also included from a study conducted by the Johns Hopkins University Applied Physics Laboratory (JHU/APL) titled Advanced Physics System Study for Future Aerospace Vehicles, which was performed under the sponsorship of the U.S. Air Force Research Laboratory Propulsion and Power Directorate and Air Vehicles Directorate. Within this study, advanced plasma technologies with application to hypersonic flight vehicles were investigated and a system-level study was conducted to demonstrate or bound the potential performance benefits of these technologies. Author
Hypersonic Vehicles; Plasmas (Physics); Technology Utilization
20060010488 Politecnico di Milano, Milan, Italy
Fundamentals of Hypersonic Flight - Properties of High Temperature Gases
Barbante, P. F.; Magin, T. E.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 5-1 - 5-50; In English; See also 20060010486; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
It is well known that, when an aerospace vehicle travels at high speed through the atmosphere, a strong shock is formed in front of it. A major part of the kinetic energy of the free stream flow is converted into thermal energy across the shock and therefore high temperature is reached in the flow region between the shock and the body (the shock layer). The intense friction happening in the boundary layer increases too the temperature triggering further chemical reactions. The high temperature effects can have a strong impact on boundary layer stability and transition to turbulence. When the shock layer temperature is high enough the gas can ionize: the free electrons absorb radio waves and cause communication blackout to and from the vehicle. This is a serious problem and an accurate prediction of the electron number density in the shock layer is important. Emission and absorption of radiation can occur and, besides affecting the state of the gas surrounding the vehicle, can raise the heat flux experienced by the vehicle itself. Radiation from the hot vehicle wall to the ambient atmosphere can have a significant cooling effect and must be taken into account in the thermal boundary condition (Sarma, 2000). Derived from text
Temperature Effects; High Temperature Gases; Hypersonic Flight; Electron Density (Concentration); Boundary Layer Stability; Turbulence; Thermal Energy; Shock Layers; Chemical Reactions
20060010490 Air Force Research Lab., Wright-Patterson AFB, OH, USA
Current and Near-Term RLV/Hypersonic Vehicle Programs
Erbland, Peter J.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 2-1 - 2-25; In English; See also 20060010486; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
This lecture provides a comprehensive review of current and near-term national and international programs aimed at developing hypersonic flight demonstration vehicles. The paper first explores the motivation for hypersonic vehicle development and potential military and civilian applications for such systems. This provides a proper framework for the discussion of various demonstration programs, which are organized according to national affiliation. Programs in the USA, Europe, the UK, and Australia are reviewed in detail and those in Japan are summarized. Where possible, context for current programs is provided through a historical perspective of recent national activities and initiatives. For each program, information is provided on the background, overall goal, key program objectives or success criteria, vehicle concept and configuration, concept of operations and/or flight test approach, critical technical challenges, key demonstrations for the flight program, development schedule/milestones and participants and their roles.
This lecture presents a survey of current and near term hypersonic vehicle demonstration programs. An effort has been made to capture most of the major national and international programs currently underway. The goal of this paper is to provide a motivation for, and context within which, technology development needs can be discussed. Many of the other lectures in this series will then focus on those individual technologies.
For the purposes of this discussion, programs will be divided into two general categories, those most closely related to development of Reusable Launch/Reentry capabilities, and those more directly supporting Hypersonic Airbreathing Propulsion. This distinction is arbitrary but useful for grouping programs of similar focus.
The scope of the 'RLV/hypersonic vehicle' classification in the title of this paper includes rocket and airbreathing reusable launch vehicle programs as well as powered, boost-glide, and un-powered hypersonic vehicle programs. History has shown hypersonic system development programs and flight demonstration programs to be very dynamic and somewhat unstable. Most programs proposed over the past 20 years have been terminated prior to reaching flight, for example; Hermes, HOTOL, Sanger, the National AeroSpace Plane (NASP) and several recent NASA X-vehicle programs including the X-33, X-34, X-38 and X-43B/C. The system development efforts have often suffered from an immature technology base and the technology demonstrators from a weak connection to future systems. Derived from text
Hypersonic Vehicles; Classifications; Hypersonic Flight; Reusable Launch Vehicles
20060010491 Von Karman Inst. for Fluid Dynamics, Rhode-Saint-Genese, Belgium
Fundamentals of Hypersonic Flow-Aerothermodynamics
Fletcher, D. G.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 3-1 - 3-47; In English; See also 20060010486; Original contains black and white illustrations; Copyright; Avail.: CASI: A04, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
The hypersonic flight regime includes atmospheric entry and re-entry, ground testing, and flight for both powered and unpowered vehicles. In the present Lecture Series, the main interest is on sustained and controlled hypersonic flight, whether for military or civil transport application.
Even though it is not currently certified for flight, there is one operational hypersonic vehicle: the space shuttle of NASA. At least 20 years before the development of the shuttle a significant activity in hypersonic flight research was conducted by the US Air Force in their X-15 program. This vehicle has reached a flight Mach number of 6.7 on its final flight, which also used to test a hypersonic ramjet engine. Direct shock impingement on the pylon holding a dummy engine caused severe heating and structural damage, and this was one of many lessons learned from the program. Owing to the design of the X-15, it was not capable of long-duration powered flight, but it provided a great deal of information on technical problems that still remain a serious obstacle to the development of new hypersonic vehicles. It is still astonishing to look back on the rapid development of high speed flight in the years after the second World War. The challenge is to build on this experience, and to accomplish the development of a new generation of flight vehicles. To put sustained hypersonic flight in context, current and proposed hypersonic vehicle trajectories are compared in Fig. 1.1.
Although un-powered hypersonic vehicles are not the topic of the present Lecture Series it is important to note that there have been many more successful developments of these types of vehicle, predominantly in the reentry of manned and unmanned spacecraft of Russian, American, and European origin into earths atmosphere. For example, the Apollo reentry conditions were 53 km altitude, 11 km/s velocity, 270 K temperature, and speed of sound 338 m/s which gave a reentry Mach number of M = 32.5.
There have also been a number of missions to other planets (more vehicles going to these planets than have been developed for sustained hypersonic flight within the atmosphere) and the entry speeds into those atmospheres have been even greater. A recent, noteworthy example of this was the Galileo probe to Jupiter that was designed to enter the Jovian atmosphere at 60 km/s at an altitude of 1000 km. At this altitude, the temperature is approximately 800 K, and the atmosphere was assumed to consist of H2 and He at a mixture of (89:11) by mass. Therefore, the entry Mach number was about 28 for this mission. Even though the entry speed was greater than that of the Apollo reentry, the Mach number is lower owing to the greater value of the sound speed in the hydrogen-helium atmosphere. Derived from text
Aerothermodynamics; Atmospheric Entry; Flight Tests; Hypersonic Flight; Hypersonic Speed; Hypersonic Vehicles; Space Shuttles; Flight Control
20060010495 Florida Univ., Gainesville, FL, USA
Propulsion Systems for Hypersonic Flight
Segal, C.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 10-1 - 10-24; In English; See also 20060010486; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
Given the broad range of aerothermodynamic conditions experienced during flight over the altitude-velocity envelope that is experienced by a scramjet-equipped vehicle it is likely that the scramjet operation will be combined with other propulsion modes. Furthermore, as the flight speed increases the vehicle aerodynamic characteristics and the engine performance become closely coupled; the vehicle forebody becomes part of the engine intake and the vehicle aft becomes part of the engine nozzle; engine throttling changes the pressure distribution on the lower part of the fuselage to a significant degree modifying the moments acting on the vehicle.
These considerations include not only the structure of the flowfield generated on the forebody of the vehicle, which is substantial given the low angle of high Mach shock waves, but also the cooling requirements which become increasingly higher with the flight speed and must be assisted by the fuel present on board. This close coupling between engine and the structure requires that, in fact, the selection of the engine cycle will be dictated by the entire system optimization.
Performance based differences between the different engine cycles are clearly illustrated in the fuel specific impulse diagram. The diagram shows that around Mach 3 flight regime the subsonic combustion ramjet becomes more efficient as a propulsive system in comparison with the turbine based engines (turbojets of turbofans) but beyond Mach 5 its specific impulse decays rapidly and the scramjet delivers a higher specific impulse at higher speed. The rocket's specific impulse is considerably lower than the other propulsion system but it offers operation capabilities from sea-level static to beyond the atmosphere which no other propulsion system mentioned here can do. The low specific impulse in comparison with the other propulsion systems clearly eliminates the rocket from consideration for long range cruise but as the Mach number continues to increase in the hypersonic regime the scramjet specific impulse approaches that of the rocket engine. Since the very high Mach numbers are expected for operation close to the edge of the atmosphere the continually decreasing air density will eventually require that the engine makes the transition to rocket operation for orbit insertion.
Historically, multiple-staged vehicles have been designed to operate with a single type of propulsion system for each stage. Stages are optimized for different altitude/Mach number regimes in the trajectory, increasing the overall system specific impulse. As an example, NASA's hypersonic aircraft demonstrators, X-43, use subsonic aircraft propulsion as the first stage followed by a second stage provided by a Pegasus (first stage) rocket with the scramjet-based research vehicle as the third stage. This limited-range accelerator begins its autonomous flight at M greater than 7. Derived from text
Hypersonic Flight; Propulsion System Configurations; Supersonic Combustion Ramjet Engines
20060010499 Technische Hochschule, Munich, Germany
Historical Perspective on Programs, Vehicles and Technology Issues
Hirschel, E. H.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 1-1 - 1-22; In English; See also 20060010486; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
System definition and technology development and preparation for hypersonic flight is an ongoing process since about 60 years. However, only one operational flight vehicle, the U. S. Space Shuttle, has resulted so far. Nevertheless, much insight and an appreciable technology base has been created, and several experimental vehicles have flown and contributed to this. In this paper programs, vehicles and technology issues are considered concerning space-trans-portation systems and hypersonic flight vehicles in general. Purely military hypersonic flight systems are not taken into account. The discussion is organised along two flight-vehicle classes: aeroassisted re-entry vehicles (RV-type vehicles) and airbreathing cruise and acceleration vehicles (CAV-type vehicles), with little consideration of shades in between. The speed regime of these vehicles is below approximately 8.0 km/s at altitudes below approximately 100.0 km. The paper is organised such, that after the classification of vehicles air-vehicle engineering issues are discussed with an overview over the requirements on the technical disciplines and a consideration of some special issues of CAV-type vehicles. Major programs and flight vehicles are then discussed, first of RV-type, then of CAV-type. Finally selected design and development issues from the HERMES and the SANGER work are treated and the implications of the second mathematisation wave in sciences and engineering for the design and development of hypersonic flight vehicles technology are discussed. Throughout the paper an air-vehicle engineering perspective is taken, technical disciplines are not considered in depth, except to a certain degree aerothermodynamics in its multidisciplinary context. Completeness is not attempted. Derived from text
Reentry Vehicles; Aeroassist; Space Shuttles; Hypersonic Vehicles; Hypersonic Flight; Classifications; Aerothermodynamics; Histories
20060010500 NASA Ames Research Center, Moffett Field, CA, USA
Shuttle Orbiter Contingency Abort Aerodynamics: Real-Gas Effects and High Angles of Attack
Prabhu, Dinesh K.; Papadopoulos, Periklis E.; Davies, Carol B.; Wright, Michael J.; McDaniel, Ryan D.; Venkatapathy, Ethiraj; Wercinski, Paul F.; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 11-1 - DP-17; In English; See also 20060010486; Original contains color and black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
An important element of the Space Shuttle Orbiter safety improvement plan is the improved understanding of its aerodynamic performance so as to minimize the 'black zones' in the contingency abort trajectories [1]. These zones are regions in the launch trajectory where it is predicted that, due to vehicle limitations, the Orbiter will be unable to return to the launch site in a two or three engine-out scenario. Reduction of these zones requires accurate knowledge of the aerodynamic forces and moments to better assess the structural capability of the vehicle. An interesting aspect of the contingency abort trajectories is that the Orbiter would need to achieve angles of attack as high as 60deg. Such steep attitudes are much higher than those for a nominal flight trajectory. The Orbiter is currently flight certified only up to an angle of attack of 44deg at high Mach numbers and has never flown at angles of attack larger than this limit. Contingency abort trajectories are generated using the data in the Space Shuttle Operational Aerodynamic Data Book (OADB) [2]. The OADB, a detailed document of the aerodynamic environment of the current Orbiter, is primarily based on wind-tunnel measurements (over a wide Mach number and angle-of-attack range) extrapolated to flight conditions using available theories and correlations, and updated with flight data where available. For nominal flight conditions, i.e., angles of attack of less than 45deg, the fidelity of the OADB is excellent due to the availability of flight data. However, at the off-nominal conditions, such as would be encountered on contingency abort trajectories, the fidelity of the OADB is less certain. The primary aims of a recent collaborative effort (completed in the year 2001) between NASA and Boeing were to determine: 1) accurate distributions of pressure and shear loads on the Orbiter at select points in the contingency abort trajectory space; and 2) integrated aerodynamic forces and moments for the entire vehicle and the control surfaces (body flap, speed brake, and elevons). The latter served the useful purpose of verification of the aerodynamic characteristics that went into the generation of the abort trajectories. Derived from text
Space Shuttle Orbiters; Abort Trajectories; Aerodynamic Characteristics; Mach Number; Angle of Attack; Control Surfaces; Real Gases; Aerodynamic Forces
20060010501 NASA Ames Research Center, Moffett Field, CA, USA
Strategies and Approaches to TPS Design
Kolodziej, Paul; Critical Technologies for Hypersonic Vehicle Development; December 2005, pp. 13-1 - 13-28; In English; See also 20060010486; Original contains black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
Thermal protection systems (TPS) insulate planetary probes and Earth re-entry vehicles from the aerothermal heating experienced during hypersonic deceleration to the planet s surface. The systems are typically designed with some additional capability to compensate for both variations in the TPS material and for uncertainties in the heating environment. This additional capability, or robustness, also provides a surge capability for operating under abnormal severe conditions for a short period of time, and for unexpected events, such as meteoroid impact damage, that would detract from the nominal performance. Strategies and approaches to developing robust designs must also minimize mass because an extra kilogram of TPS displaces one kilogram of payload. Because aircraft structures must be optimized for minimum mass, reliability-based design approaches for mechanical components exist that minimize mass. Adapting these existing approaches to TPS component design takes advantage of the extensive work, knowledge, and experience from nearly fifty years of reliability-based design of mechanical components. A Non-Dimensional Load Interference (NDLI) method for calculating the thermal reliability of TPS components is presented in this lecture and applied to several examples. A sensitivity analysis from an existing numerical simulation of a carbon phenolic TPS provides insight into the effects of the various design parameters, and is used to demonstrate how sensitivity analysis may be used with NDLI to develop reliability-based designs of TPS components. Author
Thermal Protection; Aerodynamic Heating; Aerothermodynamics; Design Analysis; Reentry Vehicles; Hypersonics
20060010899 NASA Langley Research Center, Hampton, VA, USA
Systems Analysis for a Venus Aerocapture Mission
Lockwood, Mary Kae; Starr, Brett R.; Paulson, John W., Jr.; Kontinos, Dean A.; Chen, Y. K.; Laub, Bernard; Olejniczak, Joseph; Wright, Michael J.; Takashima, Naruhisa; Justus, Carl G., et al.; March 2006; 23 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): 620.02.01.01.07.02 Report No.(s): NASA/TM-2006-214291; L-19237; Copyright; Avail.: CASI: A03, Hardcopy
Previous high level analysis has indicated that significant mass savings may be possible for planetary science missions if aerocapture is employed to place a spacecraft in orbit. In 2001 the In-Space Propulsion program identified aerocapture as one of the top three propulsion technologies for planetary exploration but that higher fidelity analysis was required to verify the favorable results and to determine if any supporting technology gaps exist that would enable or enhance aerocapture missions. A series of three studies has been conducted to assess, from an overall system point of view, the merit of using aerocapture at Titan, Neptune and Venus. These were chosen as representative of a moon with an atmosphere, an outer giant gas planet and an inner planet. The Venus mission, based on desirable science from plans for Solar System Exploration and Principal Investigator proposals, to place a spacecraft in a 300km polar orbit was examined and the details of the study are presented in this paper. Author
Aerocapture; Systems Analysis; Propulsion; Spacecraft Orbits
20060011004 Indian Space Research Organization, Bangalore, India
Journal of Spacecraft Technology, Volume 16, No. 1
January 2006; ISSN 0971-1600; 77 pp.; In English; See also 20060011005 - 20060011011; Original contains color illustrations; Copyright; Avail.: Other Sources
Topics covered include: High Solar Absorber Coating on Metallic, Semiconductor and Dielectric Surfaces for Optical Attenuation Applications in Spacecraft Systems; Orbit Model Studies for Onboard Orbit Estimation; New Studies Related to the Improved Performance in the Stepper Motor for Solar Array Drive Assembly (SADA) for Indian Remote Sensing (IRS) Satellites; Antenna Fine Pointing Mechanism (AFPM); Studies on Test and Evaluation of N+-P Junction Silicon Photo-Detectors for Space Qualification; X-ray Optics: A New Technology Development for ASTROSAT and Future Scientific Space Missions; and Fiber Optic Bus for Space Applications. Derived from text
Coating; Dielectrics; Wave Attenuation; Solar Arrays; Semiconductors (Materials); Remote Sensing; Space Missions; Solar Energy Absorbers
20060011011 Indian Space Research Organization, Bangalore, India
Fiber Optic Bus for Space Applications
Venugopal, U.; Narasimhulu, M. C.; Satyanarayana, K.; Journal of Spacecraft Technology, Volume 16, No. 1; January 2006, pp. 62-72; In English; See also 20060011004; Copyright; Avail.: Other Sources
In order to meet the stringent spacecraft performances and associated complex requirements, need for lightweight and high capacity communication system as its internal bus for command and control are increasing. The advantages of fiber optic bus system over- the conventional electrical bus system make it a suitable candidate for the above requirements. Activity carried out elsewhere in this regard is reviewed here with an aim to develop a fiber optic system, which could be used in future spacecrafts of ISRO. There are advantages and at the same time there are disadvantages in using this fiber optic bus. A proper hardness assurance program at the component level, careful electrical, optical and opt electrical designs and well thought out system protocol are essential to develop a qualified fiber optic link for space application. Author
Bus Conductors; Fiber Optics; Technology Utilization; Command and Control
Source: NASA
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