Flight performance of the TCV B-737 airplane at Montreal/Dorval International Airport, Montreal, Canada, using TRSB/MLS guidance

NASA Technical Reports Server (NTRS)

White, W. F.; Clark, L. V.

1980-01-01

The NASA terminal configured vehicle B-737 was flown in support of the world wide FAA demonstration of the time reference scanning beam microwave landing system. A summary of the flight performance of the TCV airplane during demonstration automatic approaches and landings while utilizing TRSB/MLS guidance is presented. The TRSB/MLS provided the terminal area guidance necessary for automatically flying curved, noise abatement type approaches and landings with short finals.

Software Partitioning Schemes for Advanced Simulation Computer Systems. Final Report.

ERIC Educational Resources Information Center

Clymer, S. J.

Conducted to design software partitioning techniques for use by the Air Force to partition a large flight simulator program for optimal execution on alternative configurations, this study resulted in a mathematical model which defines characteristics for an optimal partition, and a manually demonstrated partitioning algorithm design which…

Flight Simulator Evaluation of Synthetic Vision Display Concepts to Prevent Controlled Flight Into Terrain (CFIT)

NASA Technical Reports Server (NTRS)

Arthur, Jarvis J., III; Prinzel, Lawrence J., III; Kramer, Lynda J.; Parrish, Russell V.; Bailey, Randall E.

2004-01-01

In commercial aviation, over 30-percent of all fatal accidents worldwide are categorized as Controlled Flight Into Terrain (CFIT) accidents, where a fully functioning airplane is inadvertently flown into the ground. The major hypothesis for a simulation experiment conducted at NASA Langley Research Center was that a Primary Flight Display (PFD) with synthetic terrain will improve pilots ability to detect and avoid potential CFITs compared to conventional instrumentation. All display conditions, including the baseline, contained a Terrain Awareness and Warning System (TAWS) and Vertical Situation Display (VSD) enhanced Navigation Display (ND). Each pilot flew twenty-two approach departure maneuvers in Instrument Meteorological Conditions (IMC) to the terrain challenged Eagle County Regional Airport (EGE) in Colorado. For the final run, flight guidance cues were altered such that the departure path went into terrain. All pilots with a synthetic vision system (SVS) PFD (twelve of sixteen pilots) noticed and avoided the potential CFIT situation. The four pilots who flew the anomaly with the conventional baseline PFD configuration (which included a TAWS and VSD enhanced ND) had a CFIT event. Additionally, all the SVS display concepts enhanced the pilot s situational awareness, decreased workload and improved flight technical error (FTE) compared to the baseline configuration.

Flight testing the fixed-wing configuration of the Rotor Systems Research Aircraft (RSRA)

NASA Technical Reports Server (NTRS)

Hall, G. W.; Morris, P. M.

1985-01-01

The Rotor Systems Research Aircraft (RSRA) is a unique research aircraft designed to flight test advanced helicopter rotor system. Its principal flight test configuration is as a compound helicopter. The fixed wing configuration of the RSRA was primarily considered an energy fly-home mode in the event it became necessary to sever an unstable rotor system in flight. While it had always been planned to flight test the fixed wing configuration, the selection of the RSRA as the flight test bed for the X-wing rotor accelerated this schedule. This paper discusses the build-up to, and the test of, the RSRA fixed wing configuration. It is written primarily from the test pilot's perspective.

Distributed asynchronous microprocessor architectures in fault tolerant integrated flight systems

NASA Technical Reports Server (NTRS)

Dunn, W. R.

1983-01-01

The paper discusses the implementation of fault tolerant digital flight control and navigation systems for rotorcraft application. It is shown that in implementing fault tolerance at the systems level using advanced LSI/VLSI technology, aircraft physical layout and flight systems requirements tend to define a system architecture of distributed, asynchronous microprocessors in which fault tolerance can be achieved locally through hardware redundancy and/or globally through application of analytical redundancy. The effects of asynchronism on the execution of dynamic flight software is discussed. It is shown that if the asynchronous microprocessors have knowledge of time, these errors can be significantly reduced through appropiate modifications of the flight software. Finally, the papear extends previous work to show that through the combined use of time referencing and stable flight algorithms, individual microprocessors can be configured to autonomously tolerate intermittent faults.

Integrated Technology Rotor/Flight Research Rotor (ITR/FRR) concept definition

NASA Technical Reports Server (NTRS)

Howes, H. E.; Tomashofski, C. A.

1983-01-01

A program was performed to identify and evaluate a variety of candidate rotor hub configurations for the ITR/FRR. Design criteria were established for the development of the hub concepts. Five initial hub configurations were examined and two were selected for further refinement and evaluation. The selected configurations were bearingless designs. The Classic Elastic Pitch Beam (CEPB) and the Plain Elastic Pitch Beam (PEPB), both exhibit superior qualities for the criteria in the final evalution. The CEPB is favored over the PEPB primarily because it offers better capability for built in damping for stability and is judged to have a lower risk in development.

Final Phase Flight Performance and Touchdown Time Assessment of TDV in RLV-TD HEX-01 Mission

NASA Astrophysics Data System (ADS)

Yadav, Sandeep; Jayakumar, M.; Nizin, Aziya; Kesavabrahmaji, K.; Shyam Mohan, N.

2017-12-01

RLV-TD HEX-01 mission was configured as a precursor flight to actual two stages to orbit vehicle. In this mission RLV-TD was designed as a two stage vehicle for demonstrating the hypersonic flight of a winged body vehicle at Mach No. 5. One of the main objectives of this mission was to generate data for better understanding of new technologies required to design the future vehicle. In this mission, the RLV-TD vehicle was heavily instrumented to get data related to performance of different subsystems. As per the mission design, RLV-TD will land in sea after flight duration of 700 s and travelling a distance of nearly 500 km in Bay of Bengal from the launch site for a nominal trajectory. The visibility studies for telemetry data of vehicle for the nominal and off nominal trajectories were carried out. Based on that, three ground stations were proposed for the telemetry data reception (including one in sea). Even with this scheme it was seen that during the final phase of the flight there will not be any ground station visible to the flight due to low elevation. To have the mission critical data during final phase of the flight, telemetry through INSAT scheme was introduced. During the end of the mission RLV-TD will be landing in the sea on a hypothetical runway. To know the exact time of touchdown for the flight in sea, there was no direct measurement available. Simultaneously there were all chances of losing ground station visibility just before touchdown, making it difficult to assess flight performance during that phase. In this work, telemetry and instrumentation scheme of RLV-TD HEX-01 mission is discussed with an objective to determine the flight performance during the final phase. Further, using various flight sensor data the touchdown time of TDV is assessed for this mission.

Reduced gravity multibody dynamics testing

NASA Technical Reports Server (NTRS)

Sillanpaa, Meija

1993-01-01

The Final Report on reduced gravity multibody dynamics testing is presented. Tests were conducted on board the NASA KC-135 RGA in Houston, Texas. The objective was to analyze the effects of large angle rotations on flexible, multi-segmented structures. The flight experiment was conducted to provide data which will be compared to the data gathered from ground tests of the same configurations. The flight and ground tested data will be used to validate the TREETOPS software, software which models dynamic multibody systems, and other multibody codes. The flight experiment consisted of seven complete flights on board the KC-135 RGA during two one-week periods. The first period of testing was 4-9 Apr. 1993. The second period of testing was 13-18 Jun. 1993.

Flight Test Series 3: Flight Test Report

NASA Technical Reports Server (NTRS)

Marston, Mike; Sternberg, Daniel; Valkov, Steffi

2015-01-01

This document is a flight test report from the Operational perspective for Flight Test Series 3, a subpart of the Unmanned Aircraft System (UAS) Integration in the National Airspace System (NAS) project. Flight Test Series 3 testing began on June 15, 2015, and concluded on August 12, 2015. Participants included NASA Ames Research Center, NASA Armstrong Flight Research Center, NASA Glenn Research Center, NASA Langley Research center, General Atomics Aeronautical Systems, Inc., and Honeywell. Key stakeholders analyzed their System Under Test (SUT) in two distinct configurations. Configuration 1, known as Pairwise Encounters, was subdivided into two parts: 1a, involving a low-speed UAS ownship and intruder(s), and 1b, involving a high-speed surrogate ownship and intruder. Configuration 2, known as Full Mission, involved a surrogate ownship, live intruder(s), and integrated virtual traffic. Table 1 is a summary of flights for each configuration, with data collection flights highlighted in green. Section 2 and 3 of this report give an in-depth description of the flight test period, aircraft involved, flight crew, and mission team. Overall, Flight Test 3 gathered excellent data for each SUT. We attribute this successful outcome in large part from the experience that was acquired from the ACAS Xu SS flight test flown in December 2014. Configuration 1 was a tremendous success, thanks to the training, member participation, integration/testing, and in-depth analysis of the flight points. Although Configuration 2 flights were cancelled after 3 data collection flights due to various problems, the lessons learned from this will help the UAS in the NAS project move forward successfully in future flight phases.

Shuttle payload S-band communications system

NASA Technical Reports Server (NTRS)

Batson, B. H.; Teasdale, W. E.; Pawlowski, J. F.; Schmidt, O. L.

1985-01-01

The Shuttle payload S-band communications system design, operational capabilities, and performance are described in detail. System design requirements, overall system and configuration and operation, and laboratory/flight test results are presented. Payload communications requirements development is discussed in terms of evolvement of requirements as well as the resulting technical challenges encountered in meeting the initial requirements. Initial design approaches are described along with cost-saving initiatives that subsequently had to be made. The resulting system implementation that was finally adopted is presented along with a functional description of the system operation. A description of system test results, problems encountered, how the problems were solved, and the system flight experience to date is presented. Finally, a summary of the advancements made and the lessons learned is discussed.

Advanced Command Destruct System (ACDS) Enhanced Flight Termination System (EFTS)

NASA Technical Reports Server (NTRS)

Tow, David

2009-01-01

NASA Dryden started working towards a single vehicle enhanced flight termination system (EFTS) in January 2008. NASA and AFFTC combined their efforts to work towards final operating capability for multiple vehicle and multiple missions simultaneously, to be completed by the end of 2011. Initially, the system was developed to support one vehicle and one frequency per mission for unmanned aerial vehicles (UAVs) at NASA Dryden. By May 2008 95% of design and hardware builds were completed, however, NASA Dryden's change of software safety scope and requirements caused delays after May 2008. This presentation reviews the initial and final operating capabilities for the Advanced Command Destruct System (ACDS), including command controller and configuration software development. A requirements summary is also provided.

Space shuttle avionics system

NASA Technical Reports Server (NTRS)

Hanaway, John F.; Moorehead, Robert W.

1989-01-01

The Space Shuttle avionics system, which was conceived in the early 1970's and became operational in the 1980's represents a significant advancement of avionics system technology in the areas of systems and redundacy management, digital data base technology, flight software, flight control integration, digital fly-by-wire technology, crew display interface, and operational concepts. The origins and the evolution of the system are traced; the requirements, the constraints, and other factors which led to the final configuration are outlined; and the functional operation of the system is described. An overall system block diagram is included.

Software control and system configuration management - A process that works

NASA Technical Reports Server (NTRS)

Petersen, K. L.; Flores, C., Jr.

1983-01-01

A comprehensive software control and system configuration management process for flight-crucial digital control systems of advanced aircraft has been developed and refined to insure efficient flight system development and safe flight operations. Because of the highly complex interactions among the hardware, software, and system elements of state-of-the-art digital flight control system designs, a systems-wide approach to configuration control and management has been used. Specific procedures are implemented to govern discrepancy reporting and reconciliation, software and hardware change control, systems verification and validation testing, and formal documentation requirements. An active and knowledgeable configuration control board reviews and approves all flight system configuration modifications and revalidation tests. This flexible process has proved effective during the development and flight testing of several research aircraft and remotely piloted research vehicles with digital flight control systems that ranged from relatively simple to highly complex, integrated mechanizations.

Software control and system configuration management: A systems-wide approach

NASA Technical Reports Server (NTRS)

Petersen, K. L.; Flores, C., Jr.

1984-01-01

A comprehensive software control and system configuration management process for flight-crucial digital control systems of advanced aircraft has been developed and refined to insure efficient flight system development and safe flight operations. Because of the highly complex interactions among the hardware, software, and system elements of state-of-the-art digital flight control system designs, a systems-wide approach to configuration control and management has been used. Specific procedures are implemented to govern discrepancy reporting and reconciliation, software and hardware change control, systems verification and validation testing, and formal documentation requirements. An active and knowledgeable configuration control board reviews and approves all flight system configuration modifications and revalidation tests. This flexible process has proved effective during the development and flight testing of several research aircraft and remotely piloted research vehicles with digital flight control systems that ranged from relatively simple to highly complex, integrated mechanizations.

Software Considerations for Subscale Flight Testing of Experimental Control Laws

NASA Technical Reports Server (NTRS)

Murch, Austin M.; Cox, David E.; Cunningham, Kevin

2009-01-01

The NASA AirSTAR system has been designed to address the challenges associated with safe and efficient subscale flight testing of research control laws in adverse flight conditions. In this paper, software elements of this system are described, with an emphasis on components which allow for rapid prototyping and deployment of aircraft control laws. Through model-based design and automatic coding a common code-base is used for desktop analysis, piloted simulation and real-time flight control. The flight control system provides the ability to rapidly integrate and test multiple research control laws and to emulate component or sensor failures. Integrated integrity monitoring systems provide aircraft structural load protection, isolate the system from control algorithm failures, and monitor the health of telemetry streams. Finally, issues associated with software configuration management and code modularity are briefly discussed.

Research on the F/A-18E/F Using a 22%-Dynamically-Scaled Drop Model

NASA Technical Reports Server (NTRS)

Croom, M.; Kenney, H.; Murri, D.; Lawson, K.

2000-01-01

Research on the F/A-18E/F configuration was conducted using a 22%-dynamically-scaled drop model to study flight dynamics in the subsonic regime. Several topics were investigated including longitudinal response, departure/spin resistance, developed spins and recoveries, and the failing leaf mode. Comparisons to full-scale flight test results were made and show the drop model strongly correlates to the airplane even under very dynamic conditions. The capability to use the drop model to expand on the information gained from full-scale flight testing is also discussed. Finally, a preliminary analysis of an unusual inclined spinning motion, dubbed the "cartwheel", is presented here for the first time.

In space performance of the lunar orbiter laser altimeter (LOLA) laser transmitter

NASA Astrophysics Data System (ADS)

Yu, Anthony W.; Shaw, George B.; Novo-Gradac, Ann Marie; Li, Steven X.; Cavanaugh, John

2011-11-01

In this paper we present the final configuration of the space flight laser transmitter as delivered to the Lunar Orbiter Laser Altimeter (LOLA) instrument along with some in-space operation performance data. The LOLA instrument is designed to map the lunar surface and provide unprecedented data products in anticipation of future manned flight missions. The laser transmitter has been operating on orbit at the Moon continuously since July 2009 and accumulated over 1.8 billion laser shots in space. The LOLA laser transmitter design has heritage dated back to the MOLA laser transmitter launched more than 10 years ago and incorporates lessons learned from previous laser altimeter missions at NASA Goddard Space Flight Center.

Development and evaluation of automatic landing control laws for light wing loading STOL aircraft

NASA Technical Reports Server (NTRS)

Feinreich, B.; Degani, O.; Gevaert, G.

1981-01-01

Automatic flare and decrab control laws were developed for NASA's experimental Twin Otter. This light wing loading STOL aircraft was equipped with direct lift control (DLC) wing spoilers to enhance flight path control. Automatic landing control laws that made use of the spoilers were developed, evaluated in a simulation and the results compared with these obtained for configurations that did not use DLC. The spoilers produced a significant improvement in performance. A simulation that could be operated faster than real time in order to provide statistical landing data for a large number of landings over a wide spectrum of disturbances in a short time was constructed and used in the evaluation and refinement of control law configurations. A longitudinal control law that had been previously developed and evaluated in flight was also simulated and its performance compared with that of the control laws developed. Runway alignment control laws were also defined, evaluated, and refined to result in a final recommended configuration. Good landing performance, compatible with Category 3 operation into STOL runways, was obtained.

Constrained simultaneous multi-state reconfigurable wing structure configuration optimization

NASA Astrophysics Data System (ADS)

Snyder, Matthew

A reconfigurable aircraft is capable of in-flight shape change to increase mission performance or provide multi-mission capability. Reconfigurability has always been a consideration in aircraft design, from the Wright Flyer, to the F-14, and most recently the Lockheed-Martin folding wing concept. The Wright Flyer used wing-warping for roll control, the F-14 had a variable-sweep wing to improve supersonic flight capabilities, and the Lockheed-Martin folding wing demonstrated radical in-flight shape change. This dissertation will examine two questions that aircraft reconfigurability raises, especially as reconfiguration increases in complexity. First, is there an efficient method to develop a light weight structure which supports all the loads generated by each configuration? Second, can this method include the capability to propose a sub-structure topology that weighs less than other considered designs? The first question requires a method that will design and optimize multiple configurations of a reconfigurable aerostructure. Three options exist, this dissertation will show one is better than the others. Simultaneous optimization considers all configurations and their respective load cases and constraints at the same time. Another method is sequential optimization which considers each configuration of the vehicle one after the other - with the optimum design variable values from the first configuration becoming the lower bounds for subsequent configurations. This process repeats for each considered configuration and the lower bounds update as necessary. The third approach is aggregate combination — this method keeps the thickness or area of each member for the most critical configuration, the configuration that requires the largest cross-section. This research will show that simultaneous optimization produces a lower weight and different topology for the considered structures when compared to the sequential and aggregate techniques. To answer the second question, the developed optimization algorithm combines simultaneous optimization with a new method for determining the optimum location of the structural members of the sub-structure. The method proposed here considers an over-populated structural model, one in which there are initially more members than necessary. Using a unique iterative process, the optimization algorithm removes members from the design if they do not carry enough load to justify their presence. The initial set of members includes ribs, spars and a series of cross-members that diagonally connect the ribs and spars. The final result is a different structure, which is lower weight than one developed from sequential optimization or aggregate combination, and suggests the primary load paths. Chapter 1 contains background information on reconfigurable aircraft and a description of the new reconfigurable air vehicle being considered by the Air Vehicles Directorate of the Air Force Research Laboratory. This vehicle serves as a platform to test the proposed optimization process. Chapters 2 and 3 overview the optimization method and Chapter 4 provides some background analysis which is unique to this particular reconfigurable air vehicle. Chapter 5 contains the results of the optimizations and demonstrates how changing constraints or initial configuration impacts the final weight and topology of the wing structure. The final chapter contains conclusions and comments on some future work which would further enhance the effectiveness of the simultaneous reconfigurable structural topology optimization process developed and used in this dissertation.

UAS in the NAS Flight Test Series 4 Overview

NASA Technical Reports Server (NTRS)

Murphy, Jim

2016-01-01

Flight Test Series 4 (FT4) provides the researchers with an opportunity to expand on the data collected during the first flight tests. Following Flight Test Series 3, additional scripted encounters with different aircraft performance and sensors will be conducted. FT4 is presently planned for Spring of 2016 to ensure collection of data to support the validation of the final RTCA Phase 1 DAA (Detect and Avoid) Minimum Operational Performance Standards (MOPS). There are three research objectives associated with this goal: Evaluate the performance of the DAA system against cooperative and non-cooperative aircraft encounters Evaluate UAS (Unmanned Aircraft Systems) pilot performance in response to DAA maneuver guidance and alerting with live intruder encounters Evaluate TCAS/DAA (Traffic Alert and Collision Avoidance System/Detect and Avoid) interoperability. This flight test series will focus on only the Scripted Encounters configuration, supporting the collection of data to validate the interoperability of DAA and collision avoidance algorithms.

Doppler Navigation System with a Non-Stabilized Antenna as a Sea-Surface Wind Sensor.

PubMed

Nekrasov, Alexey; Khachaturian, Alena; Veremyev, Vladimir; Bogachev, Mikhail

2017-06-09

We propose a concept of the utilization of an aircraft Doppler Navigation System (DNS) as a sea-surface wind sensor complementary to its normal functionality. The DNS with an antenna, which is non-stabilized physically to the local horizontal with x -configured beams, is considered. We consider the wind measurements by the DNS configured in the multi-beam scatterometer mode for a rectilinear flight scenario. The system feasibility and the efficiency of the proposed wind algorithm retrieval are supported by computer simulations. Finally, the associated limitations of the proposed approach are considered.

KC-135 winglet program overview

NASA Technical Reports Server (NTRS)

Barber, M. R.; Selegan, D.

1982-01-01

A joint NASA/USAF program was conducted to accomplish the following objectives: (1) evaluate the benefits that could be achieved from the application of winglets to KC-135 aircraft; and (2) determine the ability of wind tunnel tests and analytical analysis to predict winglet characteristics. The program included wind-tunnel development of a test winglet configuration; analytical predictions of the changes to the aircraft resulting from the application of the test winglet; and finally, flight tests of the developed configuration. Pressure distribution, loads, stability and control, buffet, fuel mileage, and flutter data were obtained to fulfill the objectives of the program.

Doppler Navigation System with a Non-Stabilized Antenna as a Sea-Surface Wind Sensor

PubMed Central

Nekrasov, Alexey; Khachaturian, Alena; Veremyev, Vladimir; Bogachev, Mikhail

2017-01-01

We propose a concept of the utilization of an aircraft Doppler Navigation System (DNS) as a sea-surface wind sensor complementary to its normal functionality. The DNS with an antenna, which is non-stabilized physically to the local horizontal with x-configured beams, is considered. We consider the wind measurements by the DNS configured in the multi-beam scatterometer mode for a rectilinear flight scenario. The system feasibility and the efficiency of the proposed wind algorithm retrieval are supported by computer simulations. Finally, the associated limitations of the proposed approach are considered. PMID:28598374

MSFC Skylab Orbital Workshop, volume 1. [systems analysis and equipment specifications for orbital laboratory

NASA Technical Reports Server (NTRS)

1974-01-01

The technical aspects of the Skylab-Orbital Workshop are discussed. Original concepts, goals, design philosophy, hardware, and testing are reported. The final flight configuration, overall test program, and mission performance are analyzed. The systems which are examined are: (1) the structural system, (2) the meteoroid shield systems, and (3) the environmental/thermal control subsystem.

NASTRAN Modeling of Flight Test Components for UH-60A Airloads Program Test Configuration

NASA Technical Reports Server (NTRS)

Idosor, Florentino R.; Seible, Frieder

1993-01-01

Based upon the recommendations of the UH-60A Airloads Program Review Committee, work towards a NASTRAN remodeling effort has been conducted. This effort modeled and added the necessary structural/mass components to the existing UH-60A baseline NASTRAN model to reflect the addition of flight test components currently in place on the UH-60A Airloads Program Test Configuration used in NASA-Ames Research Center's Modern Technology Rotor Airloads Program. These components include necessary flight hardware such as instrument booms, movable ballast cart, equipment mounting racks, etc. Recent modeling revisions have also been included in the analyses to reflect the inclusion of new and updated primary and secondary structural components (i.e., tail rotor shaft service cover, tail rotor pylon) and improvements to the existing finite element mesh (i.e., revisions of material property estimates). Mode frequency and shape results have shown that components such as the Trimmable Ballast System baseplate and its respective payload ballast have caused a significant frequency change in a limited number of modes while only small percent changes in mode frequency are brought about with the addition of the other MTRAP flight components. With the addition of the MTRAP flight components, update of the primary and secondary structural model, and imposition of the final MTRAP weight distribution, modal results are computed representative of the 'best' model presently available.

Full-Carpet Design of a Low-Boom Demonstrator Concept

NASA Technical Reports Server (NTRS)

Ordaz, Irian; Wintzer, Mathias; Rallabhandi, Sriram K.

2015-01-01

The Cart3D adjoint-based design framework is used to mitigate the undesirable o -track sonic boom properties of a demonstrator concept designed for low-boom directly under the flight path. First, the requirements of a Cart3D design mesh are determined using a high-fidelity mesh adapted to minimize the discretization error of the CFD analysis. Low-boom equivalent area targets are then generated at the under-track and one off-track azimuthal position for the baseline configuration. The under-track target is generated using a trim- feasible low-boom target generation process, ensuring that the final design is not only low-boom, but also trimmed at the specified flight condition. The o -track equivalent area target is generated by minimizing the A-weighted loudness using an efficient adjoint-based approach. The configuration outer mold line is then parameterized and optimized to match the off-body pressure distributions prescribed by the low-boom targets. The numerical optimizer uses design gradients which are calculated using the Cart3D adjoint- based design capability. Optimization constraints are placed on the geometry to satisfy structural feasibility. The low-boom properties of the final design are verified using the adaptive meshing approach. This analysis quantifies the error associated with the CFD mesh that is used for design. Finally, an alternate mesh construction and target positioning approach offering greater computational efficiency is demonstrated and verified.

Analysis and Preliminary Design of an Advanced Technology Transport Flight Control System

NASA Technical Reports Server (NTRS)

Frazzini, R.; Vaughn, D.

1975-01-01

The analysis and preliminary design of an advanced technology transport aircraft flight control system using avionics and flight control concepts appropriate to the 1980-1985 time period are discussed. Specifically, the techniques and requirements of the flight control system were established, a number of candidate configurations were defined, and an evaluation of these configurations was performed to establish a recommended approach. Candidate configurations based on redundant integration of various sensor types, computational methods, servo actuator arrangements and data-transfer techniques were defined to the functional module and piece-part level. Life-cycle costs, for the flight control configurations, as determined in an operational environment model for 200 aircraft over a 15-year service life, were the basis of the optimum configuration selection tradeoff. The recommended system concept is a quad digital computer configuration utilizing a small microprocessor for input/output control, a hexad skewed set of conventional sensors for body rate and body acceleration, and triple integrated actuators.

Wind-tunnel development of an SR-71 aerospike rocket flight test configuration

NASA Technical Reports Server (NTRS)

Smith, Stephen C.; Shirakata, Norm; Moes, Timothy R.; Cobleigh, Brent R.; Conners, Timothy H.

1996-01-01

A flight experiment has been proposed to investigate the performance of an aerospike rocket motor installed in a lifting body configuration. An SR-71 airplane would be used to carry the aerospike configuration to the desired flight test conditions. Wind-tunnel tests were completed on a 4-percent scale SR-71 airplane with the aerospike pod mounted in various locations on the upper fuselage. Testing was accomplished using sting and blade mounts from Mach 0.6 to Mach 3.2. Initial test objectives included assessing transonic drag and supersonic lateral-directional stability and control. During these tests, flight simulations were run with wind-tunnel data to assess the acceptability of the configurations. Early testing demonstrated that the initial configuration with the aerospike pod near the SR-71 center of gravity was unsuitable because of large nosedown pitching moments at transonic speeds. The excessive trim drag resulting from accommodating this pitching moment far exceeded the excess thrust capability of the airplane. Wind-tunnel testing continued in an attempt to find a configuration suitable for flight test. Multiple configurations were tested. Results indicate that an aft-mounted model configuration possessed acceptable performance, stability, and control characteristics.

NASA HERMeS Hall Thruster Electrical Configuration Characterization

NASA Technical Reports Server (NTRS)

Peterson, Peter; Kamhawi, Hani; Huang, Wensheng; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Hofer, Richard

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This presentation will cover the electrical configuration testing of the TDU-1 HERMeS Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined are the thruster body tied to facility ground, thruster floating, and finally the thruster body electrically tied to cathode common. The TDU-1 HERMeS was configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Evaluation of a fault tolerant system for an integrated avionics sensor configuration with TSRV flight data

NASA Technical Reports Server (NTRS)

Caglayan, A. K.; Godiwala, P. M.

1985-01-01

The performance analysis results of a fault inferring nonlinear detection system (FINDS) using sensor flight data for the NASA ATOPS B-737 aircraft in a Microwave Landing System (MLS) environment is presented. First, a statistical analysis of the flight recorded sensor data was made in order to determine the characteristics of sensor inaccuracies. Next, modifications were made to the detection and decision functions in the FINDS algorithm in order to improve false alarm and failure detection performance under real modelling errors present in the flight data. Finally, the failure detection and false alarm performance of the FINDS algorithm were analyzed by injecting bias failures into fourteen sensor outputs over six repetitive runs of the five minute flight data. In general, the detection speed, failure level estimation, and false alarm performance showed a marked improvement over the previously reported simulation runs. In agreement with earlier results, detection speed was faster for filter measurement sensors soon as MLS than for filter input sensors such as flight control accelerometers.

Flight evaluation of configuration management system concepts during transition to the landing approach for a powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Franklin, J. A.; Innis, R. C.

1980-01-01

Flight experiments were conducted to evaluate two control concepts for configuration management during the transition to landing approach for a powered-lift STOL aircraft. NASA Ames' augmentor wing research aircraft was used in the program. Transitions from nominal level-flight configurations at terminal area pattern speeds were conducted along straight and curved descending flightpaths. Stabilization and command augmentation for attitude and airspeed control were used in conjunction with a three-cue flight director that presented commands for pitch, roll, and throttle controls. A prototype microwave system provided landing guidance. Results of these flight experiments indicate that these configuration management concepts permit the successful performance of transitions and approaches along curved paths by powered-lift STOL aircraft. Flight director guidance was essential to accomplish the task.

Configuration management issues and objectives for a real-time research flight test support facility

NASA Technical Reports Server (NTRS)

Yergensen, Stephen; Rhea, Donald C.

1988-01-01

Presented are some of the critical issues and objectives pertaining to configuration management for the NASA Western Aeronautical Test Range (WATR) of Ames Research Center. The primary mission of the WATR is to provide a capability for the conduct of aeronautical research flight test through real-time processing and display, tracking, and communications systems. In providing this capability, the WATR must maintain and enforce a configuration management plan which is independent of, but complimentary to, various research flight test project configuration management systems. A primary WATR objective is the continued development of generic research flight test project support capability, wherein the reliability of WATR support provided to all project users is a constant priority. Therefore, the processing of configuration change requests for specific research flight test project requirements must be evaluated within a perspective that maintains this primary objective.

Feasibility Study for Low Drag Acoustic Liners Final Report

NASA Technical Reports Server (NTRS)

Riedel, Brian; Wu, Jackie

2017-01-01

This report documents the design and structural analysis as a final deliverable for the Phase 1 contract activity. Also included is a community noise test plan, which is a key deliverable for Phase 2. Finally, a high-level estimate (Phase 3 deliverable) is provided for the work statement of Phases 2-4, which covers the build of two inlet test articles, planning and execution of a flight test with the test inlets, as well as data analysis and final documentation. The two test inlets will be compared to the production baseline inlet configuration. There is also a plan to test one of the inlets "hardwalled" using speed tape or some other similar tape to block the acoustic perforations.

Combined distance-of-flight and time-of-flight mass spectrometer

DOE Office of Scientific and Technical Information (OSTI.GOV)

Enke, Christie G; Ray, Steven J; Graham, Alexander W

2014-02-11

A combined distance-of-flight mass spectrometry (DOFMS) and time-of-flight mass spectrometry (TOFMS) instrument includes an ion source configured to produce ions having varying mass-to-charge ratios, a first detector configured to determine when each of the ions travels a predetermined distance, a second detector configured to determine how far each of the ions travels in a predetermined time, and a detector extraction region operable to direct portions of the ions either to the first detector or to the second detector.

Shuttle payload bay thermal environments: Summary and conclusion report for STS Flights 1-5

NASA Technical Reports Server (NTRS)

Fu, J. H.; Graves, G. R.

1987-01-01

The thermal data for the payload bay of the first five shuttle flights is summarized and the engineering evaluation of that data is presented. After a general discussion on mission profiles and vehicle configurations, the thermal design and flight instrumentation systems of the payload bay are described. The thermal flight data sources and a categorization of the data are then presented. A thermal flight data summarization section provides temperature data for the five phases of a typical mission profile. These are: prelaunch, ascent, on-orbit, entry and postlanding. The thermal flight data characterization section encompasses this flight data for flight to flight variations, payload effects, temperature ranges, and other variations. Discussion of the thermal environment prediction models in use by industry and various NASA Centers, and the results predicted by these models, is followed by an evaluation of the correlation between the actual flight data and the results predicted by the models. Finally, the available thermal data are evaluated from the viewpoint of the user concerned with establishing the thermal environment in the payload bay. The data deficiencies are discussed and recommendations for their elimination are presented.

Evaluation of display and control concepts for a terminal configured vehicle in final approach in a windshear environment

NASA Technical Reports Server (NTRS)

Levison, W. H.

1978-01-01

A revised treatment of nonrandom inputs was incorporated in the model. Response behavior was observed for two display configurations (a pictorial EADI presentation and a flight-director configuration requiring use of a panel-mounted airspeed indicator), two control configurations (attitude and velocity control wheel steering), and two shear environments, each of which contained a head-to-tail shear and a vertical component. In general, performance trends predicted by the model were confirmed experimentally. Experimental and analytical results both indicated superiority to the EADI display with respect to regulation of height and airspeed errors. Velocity steering allowed tighter regulation of height errors, but control parameters had little influence on airspeed regulation. Model analysis indicated that display-related differences could be ascribed to differences in the quality of speed-related information provided by the two displays.

Residual Stress Measurements After Proof and Flight: ETP-0403

NASA Technical Reports Server (NTRS)

Webster, Ronald L..

1997-01-01

The intent of this testing was to evaluate the residual stresses that occur in and around the attachment details of a case stiffener segment that has been subjected to flight/recovery followed by proof loading. Not measured in this test were stresses relieved at joint disassembly due to out-of-round and interference effects, and those released by cutting the specimens out of the case segment. The test article was lightweight case stiffener segment 1U50715, S/N L023 which was flown in the forward stiffener position on flight SRM 14A and in the aft position on flight SRM24A. Both of these flights were flown with the 3 stiffener ring configuration. Stiffener L023 had a stiffener ring installed only on the aft stub in its first flight, and it had both rings installed on its second flight. No significant post flight damage was found on either flight. Finally, the segment was used on the DM-8 static test motor in the forward position. No stiffener rings were installed. It had only one proof pressurization prior to assignment to its first use, and it was cleaned and proof tested after each flight. Thus, the segment had seen 3 proof tests, two flight pressurizations, and two low intensity water impacts prior to manufacturing for use on DM-8. On DM-8 it received one static firing pressurization in the horizontal configuration. Residual stresses at the surface and in depth were evaluated by both the x-ray diffraction and neutron beam diffraction methods. The x-ray diffraction evaluations were conducted by Technology for Energy Corporation (TEC) at their facilities in Knoxville, TN. The neutron beam evaluations were done by Atomic Energy of Canada Limited (AECL) at the Chalk River Nuclear Laboratories in Ontario. The results showed general agreement with relatively high compressive residual stresses on the surface and moderate to low subsurface tensile residual stresses.

Experimental aeroelasticity history, status and future in brief

NASA Technical Reports Server (NTRS)

Ricketts, Rodney H.

1990-01-01

NASA conducts wind tunnel experiments to determine and understand the aeroelastic characteristics of new and advanced flight vehicles, including fixed-wing, rotary-wing and space-launch configurations. Review and assessments are made of the state-of-the-art in experimental aeroelasticity regarding available facilities, measurement techniques, and other means and devices useful in testing. In addition, some past experimental programs are described which assisted in the development of new technology, validated new analysis codes, or provided needed information for clearing flight envelopes of unwanted aeroelastic response. Finally, needs and requirements for advances and improvements in testing capabilities for future experimental research and development programs are described.

Configuration management issues and objectives for a real-time research flight test support facility

NASA Technical Reports Server (NTRS)

Yergensen, Stephen; Rhea, Donald C.

1988-01-01

An account is given of configuration management activities for the Western Aeronautical Test Range (WATR) at NASA-Ames, whose primary function is the conduct of aeronautical research flight testing through real-time processing and display, tracking, and communications systems. The processing of WATR configuration change requests for specific research flight test projects must be conducted in such a way as to refrain from compromising the reliability of WATR support to all project users. Configuration management's scope ranges from mission planning to operations monitoring and performance trend analysis.

Review of operational aspects of initial experiments utilizing the U.S. MLS. [microwave landing system effectiveness

NASA Technical Reports Server (NTRS)

Walsh, T. M.; Morello, S. A.; Reeder, J. P.

1976-01-01

An exercise to support the Federal Aviation Administration in demonstrating the U.S. candidate for an international microwave landing system (MLS) was conducted by NASA. During this demonstration the MLS was utilized to provide the TCV Boeing 737 research airplane with guidance for automatic control during transition from conventional RNAV to MLS RNAV in curved, descending flight; flare; touchdown; and roll-out. Flight profiles, system configuration, displays, and operating procedures used in the demonstration are described, and preliminary results of flight data analysis are discussed. Recent experiences with manually controlled flight in the NAFEC MLS environment are also discussed. The demonstration shows that in automatic three-dimensional flight, the volumetric signal coverage of the MLS can be exploited to enable a commercial carrier class airplane to perform complex curved, descending paths with precision turns into short final approaches terminating in landing and roll-out, even when subjected to strong and gusty tail and cross wind components and severe wind shear.

Recent Dynamic Measurements and Considerations for Aerodynamic Modeling of Fighter Airplane Configurations

NASA Technical Reports Server (NTRS)

Brandon, Jay M.; Foster, John V.

1998-01-01

As airplane designs have trended toward the expansion of flight envelopes into the high angle of attack and high angular rate regimes, concerns regarding modeling the complex unsteady aerodynamics for simulation have arisen. Most current modeling methods still rely on traditional body axis damping coefficients that are measured using techniques which were intended for relatively benign flight conditions. This paper presents recent wind tunnel results obtained during large-amplitude pitch, roll and yaw testing of several fighter airplane configurations. A review of the similitude requirements for applying sub-scale test results to full-scale conditions is presented. Data is then shown to be a strong function of Strouhal number - both the traditional damping terms, but also the associated static stability terms. Additionally, large effects of sideslip are seen in the damping parameter that should be included in simulation math models. Finally, an example of the inclusion of frequency effects on the data in a simulation is shown.

Efficient placement of structural dynamics sensors on the space station

NASA Technical Reports Server (NTRS)

Lepanto, Janet A.; Shepard, G. Dudley

1987-01-01

System identification of the space station dynamic model will require flight data from a finite number of judiciously placed sensors on it. The placement of structural dynamics sensors on the space station is a particularly challenging problem because the station will not be deployed in a single mission. Given that the build-up sequence and the final configuration for the space station are currently undetermined, a procedure for sensor placement was developed using the assembly flights 1 to 7 of the rephased dual keel space station as an example. The procedure presented approaches the problem of placing the sensors from an engineering, as opposed to a mathematical, point of view. In addition to locating a finite number of sensors, the procedure addresses the issues of unobserved structural modes, dominant structural modes, and the trade-offs involved in sensor placement for space station. This procedure for sensor placement will be applied to revised, and potentially more detailed, finite element models of the space station configuration and assembly sequence.

Comparison of Rolling Moment Characteristics During Roll Oscillations for a Low and a High Aspect Ratio Configuration

NASA Technical Reports Server (NTRS)

Brandon, Jay M.; Foster, John V.; Shah, Gautam H.; Gato, William; Wilborn, James E.

2004-01-01

Improvements in testing and modeling of nonlinear and unsteady aerodynamic effects for flight dynamics predictions of vehicle performance is critical to enable the design and implementation of new, innovative vehicle concepts. Any configuration which exhibits significant flow separation, nonlinear aerodynamics, control interactions or attempts maneuvering through one or more conditions such as these is, at present, a challenge to test, model or predict flight dynamic responses prior to flight. Even in flight test experiments, adequate models are not available to study and characterize the complex nonlinear and time-dependent flow effects occurring during portions of the maneuvering envelope. Traditionally, airplane designs have been conducted to avoid these areas of the flight envelope. Better understanding and characterization of these flight regimes may not only reduce risk and cost of flight test development programs, but also may pave the way for exploitation of those characteristics that increase airplane capabilities. One of the hurdles is that the nonlinear/unsteady effects appear to be configuration dependent. This paper compares some of the dynamic aerodynamic stability characteristics of two very different configurations - representative of a fighter and a transport airplane - during dynamic body-axis roll wind tunnel tests. The fighter model shows significant effects of oscillation frequency which are not as apparent for the transport configuration.

JWST's near infrared spectrograph status and performance overview

NASA Astrophysics Data System (ADS)

Te Plate, Maurice; Birkmann, Stephan; Sirianni, Marco; Rumler, Peter; Jensen, Peter; Ehrenwinkler, Ralf; Mosner, Peter; Karl, Hermann; Rapp, Robert; Wright, Ray; Wu, Rai

2016-09-01

The James Webb Space Telescope (JWST) Observatory is the follow-on mission to the Hubble Space Telescope (HST). JWST will be the biggest space telescope ever built and it will lead to astounding scientific breakthroughs. The mission will be launched in October 2018 from Kourou, French Guyana by an ESA provided Ariane 5 rocket. NIRSpec, one of the four instruments on board of the mission, recently underwent a major upgrade. New infrared detectors were installed and the Micro Shutter Assembly (MSA) was replaced as well. The rework was necessary because both systems were found to be degrading beyond a level that could be accepted. Now in its final flight configuration, NIRSpec underwent a final cryogenic performance test at NASA's Goddard Space Flight Center (GSFC) as part of the Integrated Science Instrument Module (ISIM). This paper will present a status overview and results of the recent test campaigns.

Flight demonstrations of curved, descending approaches and automatic landings using time referenced scanning beam guidance

NASA Technical Reports Server (NTRS)

White, W. F. (Compiler)

1978-01-01

The Terminal Configured Vehicle (TCV) program operates a Boeing 737 modified to include a second cockpit and a large amount of experimental navigation, guidance and control equipment for research on advanced avionics systems. Demonstration flights to include curved approaches and automatic landings were tracked by a phototheodolite system. For 50 approaches during the demonstration flights, the following results were obtained: the navigation system, using TRSB guidance, delivered the aircraft onto the 3 nautical mile final approach leg with an average overshoot of 25 feet past centerline, subjet to a 2-sigma dispersion of 90 feet. Lateral tracking data showed a mean error of 4.6 feet left of centerline at the category 1 decision height (200 feet) and 2.7 feet left of centerline at the category 2 decision height (100 feet). These values were subject to a sigma dispersion of about 10 feet. Finally, the glidepath tracking errors were 2.5 feet and 3.0 feet high at the category 1 and 2 decision heights, respectively, with a 2 sigma value of 6 feet.

Static Wind-Tunnel and Radio-Controlled Flight Test Investigation of a Remotely Piloted Vehicle Having a Delta Wing Planform

NASA Technical Reports Server (NTRS)

Yip, Long P.; Fratello, David J.; Robelen, David B.; Makowiec, George M.

1990-01-01

At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the Flight Dynamics Branch at the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted in the Langley 12 foot Low Speed Wind Tunnel to identify and improve the stability and control characteristics of the vehicle. The wind tunnel test resulted in several configuration modifications which included increased elevator size, increased vertical tail size and tail moment arm, increased rudder size and aileron size, the addition of vertical wing tip fins, and the addition of leading-edge droops on the outboard wing panel to improve stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.

J-2X Gas Generator Development Testing at NASA Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Reynolds, D. C.; Hormonzian, Carlo

2010-01-01

NASA is developing a liquid oxygen/liquid hydrogen rocket engine for upper stage and trans-lunar applications of the Ares vehicles for the Constellation program. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. Development was contracted to Pratt & Whitney Rocketdyne in 2006. Over the past several years, two phases of testing have been completed on the development of the gas generator for the J-2X engine. The hardware has progressed through a variety of workhorse injector, chamber, and feed system configurations. Several of these configurations have resulted in combustion instability of the gas generator assembly. Development of the final configuration of workhorse hardware (which will ultimately be used to verify critical requirements on a component level) has required a balance between changes in the injector and chamber hardware in order to successfully mitigate the combustion instability without sacrificing other engine system requirements. This paper provides an overview of the two completed test series, performed at NASA s Marshall Space Flight Center. The requirements, facility setup, hardware configurations, and test series progression are detailed. Significant levels of analysis have been performed in order to provide design solutions to mitigate the combustion stability issues, and these are briefly covered. Also discussed are the results of analyses related to either anomalous readings or off-nominal testing throughout the two test series.

Operational Overview for UAS Integration in the NAS Project Flight Test Series 3

NASA Technical Reports Server (NTRS)

Valkov, Steffi B.; Sternberg, Daniel; Marston, Michael

2017-01-01

The National Aeronautics and Space Administration Unmanned Aircraft Systems Integration in the National Airspace System Project has conducted a series of flight tests intended to support the reduction of barriers that prevent unmanned aircraft from flying without the required waivers from the Federal Aviation Administration. The 2015 Flight Test Series 3, supported two separate test configurations. The first configuration investigated the timing of Detect and Avoid alerting thresholds using a radar equipped unmanned vehicle and multiple live intruders flown at varying encounter geometries. The second configuration included a surrogate unmanned vehicle (flown from a ground control station, with a safety pilot on board) flying a mission in a virtual air traffic control airspace sector using research pilot displays and Detect and Avoid advisories to maintain separation from live and virtual aircraft. The test was conducted over an eight-week span within the R-2508 Special Use Airspace. Over 200 encounters were flown for the first configuration, and although the second configuration was cancelled after three data collection flights, Flight Test 3 proved to be invaluable for the purposes of planning, managing, and execution of this type of integrated flight test.

A Flight Study of the Conversion Maneuver of a Tilt-Duct VTOL Aircraft

NASA Technical Reports Server (NTRS)

Tapscott, Robert J.; Kelley, Henry L.

1960-01-01

Flight records are presented from an early flight test of a wing-tip mounted tilting-ducted-fan, vertical-take-off and landing (VTOL) aircraft configuration. Time histories of the aircraft motions, control positions, and duct pitching-moment variation are presented to illustrate the characteristics of the aircraft in hovering, in conversion from hovering to forward flight, and in conversion from forward flight to hovering. The results indicate that during essentially continuous slow level- flight conversions, this aircraft experiences excessive longitudinal trim changes. Studies have shown that the large trim changes are caused primarily by the variation of aerodynamic moments acting on the duct units. Action of the duct-induced downwash on the horizontal stabilizer during the conversion also contributes to the longitudinal trim variations. Time histories of hovering and slow vertical descent in the final stages of landing in calm air show angular motions of the aircraft as great as +/- 10 deg. about all axes. Stick and pedal displacements required to control the aircraft during the landing maneuver were on the order of 50 to 60 percent of the total travel available.

Final definition and preliminary design study for the initial atmospheric cloud physics laboratory, a Spacelab mission payload

NASA Technical Reports Server (NTRS)

1976-01-01

The following areas related to the final definition and preliminary design study of the initial atmospheric cloud physics laboratory (ACPL) were covered: (1) proposal organization, personnel, schedule, and project management, (2) proposed configurations, (3) study objectives, (4) ACPL experiment program listing and description, (5) mission/flight flexibility and modularity/commonality, (6) study plan, and (7) description of following tasks: requirement analysis and definition task flow, systems analysis and trade studies, subsystem analysis and trade studies, specifications and interface control documents, preliminary design task flow, work breakdown structure, programmatic analysis and planning, and project costs. Finally, an overview of the scientific requirements was presented.

In-flight lift and drag measurements on a first generation jet transport equipped with winglets

NASA Technical Reports Server (NTRS)

Lux, D. P.

1982-01-01

A KC-135A aircraft equipped with wing tip winglets was flight tested to demonstrate and validate the potential performance gain of the winglet concept as predicted from analytical and wind tunnel data. Flight data were obtained at cruise conditions for Mach numbers of 0.70, 0.75, and 0.80 at a nominal altitude of 36,000 ft. and winglet configurations of 15 deg cant/-4 deg incidence, 0 deg cant/-4 deg incidence, and baseline. For the Mach numbers tested the data show that the addition of winglets did not affect the lifting characteristics of the wing. However, both winglet configurations showed a drag reduction over the baseline configuration, with the best winglet configuration being the 15 deg cant/-4 deg incidence configuration. This drag reduction due to winglets also increased with increasing lift coefficient. It was also shown that a small difference exists between the 15 deg cant/-4 deg incidence flight and wind tunnel predicted data. This difference was attributed to the pillowing of the winglet skins in flight which would decrease the winglet performance.

A shadowgraph study of the National Launch System's 1 1/2 stage vehicle configuration and Heavy Lift Launch Vehicle configuration. [Using the Marshall Space Flight Center's 14-Inch Trisonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pokora, Darlene C.; Springer, Anthony M.

1994-01-01

A shadowgraph study of the National Launch System's (NLS's) 1 1/2 stage and heavy lift launch vehicle (HLLV) configurations is presented. Shadowgraphs are shown for the range of Mach numbers from Mach 0.6 to 5.0 at various angles-of-attack and roll angles. Since the 1 1/2 stage configuration is generally symmetric, no shadowgraphs of any roll angle are shown for this configuration. The major flow field phenomena over the NLS 1 1/2 stage and HLLV configurations are shown in the shadowgraphs. These shadowgraphs are used in the aerothermodynamic analysis of the external flow conditions the launch vehicle would encounter during the ascent stage of flight. The shadowgraphs presented in this study were obtained from configurations tested in the Marshall Space Flight Center's 14-Inch Trisonic Wind Tunnel during 1992.

IAE - Inflatable Antenna Experiment

NASA Image and Video Library

1996-05-20

STS077-150-022 (20 May 1996) --- After leaving the cargo bay of the Space Shuttle Endeavour, the Spartan 207/Inflatable Antenna Experiment (IAE) payload goes through the final stages its inflation process, backdropped over clouds and blue water. The view was photographed with a large format still camera on the first full day of in-space operations by the six-member crew. Managed by Goddard Space Flight Center (GSFC), Spartan is designed to provide short-duration, free-flight opportunities for a variety of scientific studies. The Spartan configuration on this flight is unique in that the IAE is part of an additional separate unit which is ejected once the experiment is completed. The IAE experiment will lay the groundwork for future technology development in inflatable space structures, which will be launched and then inflated like a balloon on-orbit.

Design verification and fabrication of active control systems for the DAST ARW-2 high aspect ratio wing, part 1

NASA Technical Reports Server (NTRS)

Mcgehee, C. R.

1986-01-01

A study was conducted under Drones for Aerodynamic and Structural Testing (DAST) program to accomplish the final design and hardware fabrication for four active control systems compatible with and ready for installation in the NASA Aeroelastic Research Wing No. 2 (ARW-2) and Firebee II drone flight test vehicle. The wing structure was designed so that Active Control Systems (ACS) are required in the normal flight envelope by integrating control system design with aerodynamics and structure technologies. The DAST ARW-2 configuration uses flutter suppression, relaxed static stability, and gust and maneuver load alleviation ACS systems, and an automatic flight control system. Performance goals and criteria were applied to individual systems and the systems collectively to assure that vehicle stability margins, flutter margins, flying qualities and load reductions are achieved.

Flight representative positive isolation disconnect

NASA Technical Reports Server (NTRS)

Rosener, A. A.; Jonkoniec, T. G.

1977-01-01

Resolutions were developed for each problem encountered and a tradeoff analysis was performed to select a final configuration for a flight representative PID (Positive Isolation Disconnect) that is reduced in size and comparable in weight and pressure drop to the developmental PID. A 6.35 mm (1/4-inch) line size PID was fabricated and tested. The flight representative PID consists of two coupled disconnect halves, each capable of fluid isolation with essentially zero clearance between them for zero leakage upon disconnect half disengagement. An interlocking foolproofing technique prevents uncoupling of disconnect halves prior to fluid isolation. Future development efforts for the Space Shuttle subsystems that would benefit from the use of the positive isolation disconnect are also recommended. Customary units were utilized for principal measurements and calculations with conversion factors being inserted in equations to convert the results to the international system of units.

Applied Meteorology Unit (AMU) Quarterly Report Fourth Quarter FY-14

NASA Technical Reports Server (NTRS)

Bauman, William H.; Crawford, Winifred C.; Watson, Leela R.; Shafer, Jaclyn

2014-01-01

Ms. Crawford completed the final report for the dual-Doppler wind field task. Dr. Bauman completed transitioning the 915-MHz and 50-MHz Doppler Radar Wind Profiler (DRWP) splicing algorithm developed at Marshall Space Flight Center (MSFC) into the AMU Upper Winds Tool. Dr. Watson completed work to assimilate data into model configurations for Wallops Flight Facility (WFF) and Kennedy Space Center/Cape Canaveral Air Force Station (KSC/CCAFS). Ms. Shafer began evaluating the a local high-resolution model she had set up previously for its ability to forecast weather elements that affect launches at KSC/CCAFS. Dr. Watson began a task to optimize the data-assimilated model she just developed to run in real time.

Design of in-flight fragment separator using high-power primary beams in the energy of a few hundred MeV/u

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kim, J. Y.; Kim, J. W.; Wan, W.

2015-07-15

An in-flight fragment separator usually requires large acceptance and high momentum resolution to minimize the loss of a rare isotope beam of interest produced at a thin target, which is especially important when {sup 238}U fission reactions in the energy of 200 MeV/u are used. The production target and beam dump are located in the pre-separator, where a beam power of up to 400 kW is dissipated. The area is surrounded by thick radiation shielding walls, which result in long drift spaces between adjacent magnetic components at various locations and an asymmetrical layout. Efforts have been made to minimize non-linearmore » effects in the pre-separator beam optics with trials of different separator configurations and correction schemes using COSY Infinity and GICOSY. The main separator is configured to be mirror symmetric such that correction with hexapole and octupole coils can be more readily applied. The separator configuration was finalized to allow the facility design to proceed and the key components including superconducting magnets have been designed and prototyped. In addition, the separator design has been evaluated using LISE++ including a set of wedge degraders at dispersive focal planes to improve the yield and purity of selected isotope beam.« less

Aerodynamic Interference Due to MSL Reaction Control System

NASA Technical Reports Server (NTRS)

Dyakonov, Artem A.; Schoenenberger, Mark; Scallion, William I.; VanNorman, John W.; Novak, Luke A.; Tang, Chun Y.

2009-01-01

An investigation of effectiveness of the reaction control system (RCS) of Mars Science Laboratory (MSL) entry capsule during atmospheric flight has been conducted. The reason for the investigation is that MSL is designed to fly a lifting actively guided entry with hypersonic bank maneuvers, therefore an understanding of RCS effectiveness is required. In the course of the study several jet configurations were evaluated using Langley Aerothermal Upwind Relaxation Algorithm (LAURA) code, Data Parallel Line Relaxation (DPLR) code, Fully Unstructured 3D (FUN3D) code and an Overset Grid Flowsolver (OVERFLOW) code. Computations indicated that some of the proposed configurations might induce aero-RCS interactions, sufficient to impede and even overwhelm the intended control torques. It was found that the maximum potential for aero-RCS interference exists around peak dynamic pressure along the trajectory. Present analysis largely relies on computational methods. Ground testing, flight data and computational analyses are required to fully understand the problem. At the time of this writing some experimental work spanning range of Mach number 2.5 through 4.5 has been completed and used to establish preliminary levels of confidence for computations. As a result of the present work a final RCS configuration has been designed such as to minimize aero-interference effects and it is a design baseline for MSL entry capsule.

Online Learning Flight Control for Intelligent Flight Control Systems (IFCS)

NASA Technical Reports Server (NTRS)

Niewoehner, Kevin R.; Carter, John (Technical Monitor)

2001-01-01

The research accomplishments for the cooperative agreement 'Online Learning Flight Control for Intelligent Flight Control Systems (IFCS)' include the following: (1) previous IFC program data collection and analysis; (2) IFC program support site (configured IFC systems support network, configured Tornado/VxWorks OS development system, made Configuration and Documentation Management Systems Internet accessible); (3) Airborne Research Test Systems (ARTS) II Hardware (developed hardware requirements specification, developing environmental testing requirements, hardware design, and hardware design development); (4) ARTS II software development laboratory unit (procurement of lab style hardware, configured lab style hardware, and designed interface module equivalent to ARTS II faceplate); (5) program support documentation (developed software development plan, configuration management plan, and software verification and validation plan); (6) LWR algorithm analysis (performed timing and profiling on algorithm); (7) pre-trained neural network analysis; (8) Dynamic Cell Structures (DCS) Neural Network Analysis (performing timing and profiling on algorithm); and (9) conducted technical interchange and quarterly meetings to define IFC research goals.

James Webb Space Telescope (JWST) Integrated Science Instruments Module (ISIM) Cryo-Vacuum (CV) Test Campaign Summary

NASA Technical Reports Server (NTRS)

Yew, Calinda; Whitehouse, Paul; Lui, Yan; Banks, Kimberly

2016-01-01

JWST Integrated Science Instruments Module (ISIM) has completed its system-level testing program at the NASA Goddard Space Flight Center (GSFC). In March 2016, ISIM was successfully delivered for integration with the Optical Telescope Element (OTE) after the successful verification of the system through a series of three cryo-vacuum (CV) tests. The first test served as a risk reduction test; the second test provided the initial verification of the fully-integrated flight instruments; and the third test verified the system in its final flight configuration. The complexity of the mission has generated challenging requirements that demand highly reliable system performance and capabilities from the Space Environment Simulator (SES) vacuum chamber. As JWST progressed through its CV testing campaign, deficiencies in the test configuration and support equipment were uncovered from one test to the next. Subsequent upgrades and modifications were implemented to improve the facility support capabilities required to achieve test requirements. This paper: (1) provides an overview of the integrated mechanical and thermal facility systems required to achieve the objectives of JWST ISIM testing, (2) compares the overall facility performance and instrumentation results from the three ISIM CV tests, and (3) summarizes lessons learned from the ISIM testing campaign.

Apollo 16 mission report. Supplement 2: Service Propulsion system final flight evaluation

NASA Technical Reports Server (NTRS)

Smith, R. J.; Wood, S. C.

1974-01-01

The Apollo 16 Mission was the sixteenth in a series of flights using Apollo flight hardware and included the fifth lunar landing of the Apollo Program. The Apollo 16 Mission utilized CSM 113 which was equipped with SPS Engine S/N 66 (Injector S/N 137). The engine configuration and expected performance characteristics are presented. Since previous flight results of the SPS have consistently shown the existence of a negative mixture ratio shift, SPS Engine S/N 66 was reorificed to increase the mixture ratio for this mission. The propellant unbalance for the two major engine firings is compared with the predicted unbalance. Although the unbalance at the end of the TEI burn is significantly different than the predicted unbalance, the propellant mixture ratio was well within limits. The SPS performed six burns during the mission, with a total burn duration of 575.3 seconds. The ignition time, burn duration and velocity gain for each of the six SPS burns are reported.

Configuration Aerodynamics: Past - Present - Future

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Agrawal, Shreekant; Bencze, Daniel P.; Kulfan, Robert M.; Wilson, Douglas L.

1999-01-01

The Configuration Aerodynamics (CA) element of the High Speed Research (HSR) program is managed by a joint NASA and Industry team, referred to as the Technology Integration Development (ITD) team. This team is responsible for the development of a broad range of technologies for improved aerodynamic performance and stability and control characteristics at subsonic to supersonic flight conditions. These objectives are pursued through the aggressive use of advanced experimental test techniques and state of the art computational methods. As the HSR program matures and transitions into the next phase the objectives of the Configuration Aerodynamics ITD are being refined to address the drag reduction needs and stability and control requirements of High Speed Civil Transport (HSCT) aircraft. In addition, the experimental and computational tools are being refined and improved to meet these challenges. The presentation will review the work performed within the Configuration Aerodynamics element in 1994 and 1995 and then discuss the plans for the 1996-1998 time period. The final portion of the presentation will review several observations of the HSR program and the design activity within Configuration Aerodynamics.

Quiet engine program flight engine design study

NASA Technical Reports Server (NTRS)

Klapproth, J. F.; Neitzel, R. E.; Seeley, C. T.

1974-01-01

The results are presented of a preliminary flight engine design study based on the Quiet Engine Program high-bypass, low-noise turbofan engines. Engine configurations, weight, noise characteristics, and performance over a range of flight conditions typical of a subsonic transport aircraft were considered. High and low tip speed engines in various acoustically treated nacelle configurations were included.

Exhaust emissions reduction for intermittent combustion aircraft engines

NASA Technical Reports Server (NTRS)

Rezy, B. J.; Stuckas, K. J.; Tucker, J. R.; Meyers, J. E.

1982-01-01

Three concepts which, to an aircraft piston engine, provide reductions in exhaust emissions of hydrocarbons and carbon monoxide while simultaneously improving fuel economy. The three chosen concepts, (1) an improved fuel injection system, (2) an improved cooling cylinder head, and (3) exhaust air injection, when combined, show a synergistic relationship in achieving these goals. In addition, the benefits of variable ignition timing were explored and both dynamometer and flight testing of the final engine configuration were accomplished.

Universal Reconfigurable Translator Module (URTM) Final Report

NASA Technical Reports Server (NTRS)

Leventhal, Edward; Machan, Roman; Jones, Rob

2009-01-01

This report describes the Universal Reconfigurable Translation Module, or URTM. The URTM was developed by Sigma Space Corporation for NASA in order to translate specific serial protocols, both logically and physically. At present, the prototype configuration has targeted MIL-STD-1553B (RT and BC), IEEE 1394b (Firewire), and ECSS-E-50-12A (SpaceWire). The objectives of this program were to study the feasibility of a configurable URTM to translate serial link data as might be used in a space-flight mission and to design, develop, document, and deliver an engineering prototype model of the URTM with a path to spaceflight. By simply connecting two of the three Physical Interface Modules (PIM) on either end of the RPTM (Reconfigurable Protocol Translator Module), the URTM then self configures via a library of interface translation functions, thereby allowing the two data links to communicate seamlessly.

Aircraft Configured for Flight in an Atmosphere Having Low Density

NASA Technical Reports Server (NTRS)

Teter, Jr., John E. (Inventor); Croom, Mark A. (Inventor); Smith, Stephen C. (Inventor); Gelhausen, Paul A. (Inventor); Hunter, Craig A. (Inventor); Riddick, Steven E. (Inventor); Guynn, Mark D. (Inventor); Paddock, David A. (Inventor)

2012-01-01

An aircraft is configured for flight in an atmosphere having a low density. The aircraft includes a fuselage, a pair of wings, and a rear stabilizer. The pair of wings extends from the fuselage in opposition to one another. The rear stabilizer extends from the fuselage in spaced relationship to the pair of wings. The fuselage, the wings, and the rear stabilizer each present an upper surface opposing a lower surface. The upper and lower surfaces have X, Y, and Z coordinates that are configured for flight in an atmosphere having low density.

Final STS-35 Columbia descent BET products and results for LaRC OEX investigations

NASA Technical Reports Server (NTRS)

Oakes, Kevin F.; Findlay, John T.; Jasinski, Rachel A.; Wood, James S.

1991-01-01

Final STS-35 'Columbia' descent Best Estimate Trajectory (BET) products have been developed for Langley Research Center (LaRC) Orbiter Experiments (OEX) investigations. Included are the reconstructed inertial trajectory profile; the Extended BET, which combines the inertial data and, in this instance, the National Weather Service atmospheric information obtained via Johnson Space Center; and the Aerodynamic BET. The inertial BET utilized Inertial Measurement Unit 1 (IMU1) dynamic measurements for deterministic propagation during the ENTREE estimation process. The final estimate was based on the considerable ground based C-band tracking coverage available as well as Tracking Data and Relay Satellite System (TDRSS) Doppler data, a unique use of the latter for endo-atmospheric flight determinations. The actual estimate required simultaneous solutions for the spacecraft position and velocity, spacecraft attitude, and six IMU parameters - three gyro biases and three accelerometer scale factor correction terms. The anchor epoch for this analysis was 19,200 Greenwich Mean Time (GMT) seconds which corresponds to an initial Shuttle altitude of approximately 513 kft. The atmospheric data incorporated were evaluated based on Shuttle derived considerations as well as comparisons with other models. The AEROBET was developed based on the Extended BET, the measured spacecraft configuration information, final mass properties, and the final Orbiter preoperation databook. The latter was updated based on aerodynamic consensus incrementals derived by the latest published FAD. The rectified predictions were compared versus the flight computed values and the resultant differences were correlated versus ensemble results for twenty-two previous STS entry flights.

Transition Analysis for the HIFiRE-5 Vehicle

NASA Technical Reports Server (NTRS)

Choudhari, Meelan M.; Chang, Chau-Lyan; Li, Fei; Berger, Karen T.; Candler, Graham V.; Kimmel, Roger

2009-01-01

The Hypersonic International Flight Research and Experimentation (HIFiRE) 5 flight experiment by Air Force Research Laboratories and Australian Defense Science and Technology Organization is designed to provide in-flight boundary-layer transition data for a canonical 3D configuration at hypersonic Mach numbers. This paper outlines the progress, to date, on boundary layer stability analysis for the HIFiRE-5 flight configuration, as well as for selected test conditions from the wind tunnel experiments supporting the flight test. At flow conditions corresponding to the end of the test window, rather large values of linear amplification factor are predicted for both second mode (N>40) and crossflow (N>20) instabilities, strongly supporting the feasibility of first in-flight measurements of natural transition on a fully three-dimensional hypersonic configuration. Additional results highlight the rich mixture of instability mechanisms relevant to a large segment of the flight trajectory, as well as the effects of angle of attack and yaw angle on the predicted transition fronts for ground facility experiments at Mach 6.

Proposed definition of the term en route in en route aircraft noise

NASA Technical Reports Server (NTRS)

Garbell, Maurice A.

1990-01-01

The problem of en route aircraft noise was examined in a formal, dedicated, setting. Whereas the general meaning of the term en route might be intuitively understood, it is suggested that a precise formal definition of the term en route would be opportune from the outset, especially since the scientific and technical investigation of the problem of noise immissions on the ground from aircraft in flight away from the airspace of an airport may conceivably lead to administrative, regulatory, and legal consequences that would mandatorily require a precise definition of the term en route. Five flight segments, with their differing airframe configurations, engine thrusts, and airspeed management, should form the basis for the differential consideration of the noise immissions perceived on the ground underneath or near the defined segments of the flightpath in en route flight, from the end of the initial climb from an airport after takeoff until the final approach to an airport.

In-flight investigation of the effects of pilot location and control system design on airplane flying qualities for approach and landing

NASA Technical Reports Server (NTRS)

Weingarten, N. C.; Chalk, C. R.

1982-01-01

The handling qualities of large airplanes in the approach and landing flight phase were studied. The primary variables were relative pilot position with respect to center of rotation, command path time delays and phase shifts, augmentation schemes and levels of augmentation. It is indicated that the approach and landing task with large airplanes is a low bandwidth task. Low equivalent short period frequencies and relatively long time delays are tolerated only when the pilot is located at considerable distance forward of the center of rotation. The control problem experienced by the pilots, when seated behind the center of rotation, tended to occur at low altitude when they were using visual cues of rate of sink and altitude. A direct lift controller improved final flight path control of the shuttle like configurations.

High Alpha Technology Program (HATP) ground test to flight comparisons

NASA Technical Reports Server (NTRS)

Hall, R. M.; Banks, D. W.; Fisher, David F.; Ghaffari, F.; Murri, D. G.; Ross, J. C.; Lanser, Wendy R.

1994-01-01

This status paper reviews the experimental ground test program of the High Alpha Technology Program (HATP). The reasons for conducting this ground test program had their origins during the 1970's when several difficulties were experienced during the development programs of both the F-18 and F-16. A careful assessment of ground test to flight correlations appeared to be important for reestablishing a high degree of confidence in our ground test methodology. The current paper will then focus on one aspect of the HATP program that is intended to improve the correlation between ground test and flight, high-alpha gritting. The importance of this work arises from the sensitivity of configurations with smooth-sided forebodies to Reynolds number. After giving examples of the effects of Reynolds number, the paper will highlight efforts at forebody gritting. Finally, the paper will conclude by summarizing the charter of the HATP Experimental Aerodynamics Working Group and future experimental testing plans.

Design verification and fabrication of active control systems for the DAST ARW-2 high aspect ratio wing. Part 2: Appendices

NASA Technical Reports Server (NTRS)

Mcgehee, C. R.

1986-01-01

This is Part 2-Appendices of a study conducted under Drones for Aerodynamic and Structural Testing (DAST) Program to accomplish the final design and hardware fabrication for four active control systems compatible with and ready for installation in the NASA Aeroelastic Research Wing No. 2 (ARW-2) and Firebee II drone flight test vehicle. The wing structure was designed so that Active Control Systems (ACS) are required in the normal flight envelope by integrating control system design with aerodynamics and structure technologies. The DAST ARW-2 configuration uses flutter suppression, relaxed static stability, and gust and maneuver load alleviation ACS systems, and an automatic flight control system. Performance goals and criteria were applied to individual systems and the systems collectively to assure that vehicle stability margins, flutter margins, flying qualities, and load reductions were achieved.

Aeroelastic character of a National Aerospace Plane demonstrator concept

NASA Technical Reports Server (NTRS)

Spain, Charles V.; Zeiler, Thomas A.; Gibbons, Michael D.; Soistmann, David L.; Pozefsky, Peter; Dejesus, Rafael O.; Brannon, Cyprian P.

1993-01-01

The paper provides an analytical assessment of the flutter character of an unclassified National Aerospace Plane configuration known as the demonstrator. Linear subsonic, supersonic, and hypersonic analysis indicate that the vehicle is prone to body-freedom flutter resulting from the decrease in vibration frequency of the all-moveable wing at high flight dynamic pressures. As the wing-pivot frequency decreases, it couples with the vehicle short-period mode resulting in dynamic instability. A similar instability sometimes occurs when the pivot mode couples with the fuselage-bending mode. Also assessed, for supersonic flight conditions, are configuration variations that include relocation of the wing further aft on the lifting-body fuselage, and the addition of body flaps to the rear of the vehicle. These changes are destabilizing because they result in severe wing-pivot/fuselage-bending instabilities at dynamic pressures lower than the instabilities indicated for the original demonstrator. Finally, a two-point wing support and actuation system concept is proposed for the National Aerospace Plane, which if developed may (according to cursory analysis) enhance overall stability.

Summary of shuttle data processing and aerodynamic performance comparisons for the first 11 flights

NASA Technical Reports Server (NTRS)

Findlay, J. T.; Kelly, G. M.; Heck, M. L.; Mcconnell, J. G.

1984-01-01

NASA Space Shuttle aerodynamic and aerothermodynamic research is but one part of the most comprehensive end-to-end flight test program ever undertaken considering: the extensive pre-flight experimental data base development; the multitude of spacecraft and remote measurements taken during entry flight; the complexity of the Orbiter aerodynamic configuration; the variety of flight conditions available across the entire speed regime; and the efforts devoted to flight data reduction throughout the aerospace community. Shuttle entry flights provide a wealth of research quality data, in essence a veritable flying wind tunnel, for use by researchers to verify and improve the operational capability of the Orbiter and provide data for evaluations of experimental facilities as well as computational methods. This final report merely summarizes the major activities conducted by the AMA, Inc. under NASA Contract NAS1-16087 as part of that interesting research. Investigators desiring more detailed information can refer to the glossary of AMA publications attached herein as Appendix A. Section I provides background discussion of software and methodology development to enable Best Estimate Trajectory (BET) generation. Actual products generated are summarized in Section II as tables which completely describe the post-flight products available from the first three-year Shuttle flight history. Summary results are presented in Section III, with longitudinal performance comparisons included as Appendices for each of the flights.

Computer Software Configuration Item-Specific Flight Software Image Transfer Script Generator

NASA Technical Reports Server (NTRS)

Bolen, Kenny; Greenlaw, Ronald

2010-01-01

A K-shell UNIX script enables the International Space Station (ISS) Flight Control Team (FCT) operators in NASA s Mission Control Center (MCC) in Houston to transfer an entire or partial computer software configuration item (CSCI) from a flight software compact disk (CD) to the onboard Portable Computer System (PCS). The tool is designed to read the content stored on a flight software CD and generate individual CSCI transfer scripts that are capable of transferring the flight software content in a given subdirectory on the CD to the scratch directory on the PCS. The flight control team can then transfer the flight software from the PCS scratch directory to the Electronically Erasable Programmable Read Only Memory (EEPROM) of an ISS Multiplexer/ Demultiplexer (MDM) via the Indirect File Transfer capability. The individual CSCI scripts and the CSCI Specific Flight Software Image Transfer Script Generator (CFITSG), when executed a second time, will remove all components from their original execution. The tool will identify errors in the transfer process and create logs of the transferred software for the purposes of configuration management.

Survey of piloting factors in V/STOL aircraft with implications for flight control system design

NASA Technical Reports Server (NTRS)

Ringland, R. F.; Craig, S. J.

1977-01-01

Flight control system design factors involved for pilot workload relief are identified. Major contributors to pilot workload include configuration management and control and aircraft stability and response qualities. A digital fly by wire stability augmentation, configuration management, and configuration control system is suggested for reduction of pilot workload during takeoff, hovering, and approach.

An Overview of the NASA F-18 High Alpha Research Vehicle

NASA Technical Reports Server (NTRS)

Bowers, Albion H.; Pahle, Joseph W.; Wilson, R. Joseph; Flick, Bradley C.; Rood, Richard L.

1996-01-01

This paper gives an overview of the NASA F-18 High Alpha Research Vehicle. The three flight phases of the program are introduced, along with the specific goals and data examples taken during each phase. The aircraft configuration and systems needed to perform the disciplinary and inter-disciplinary research are discussed. The specific disciplines involved with the flight research are introduced, including aerodynamics, controls, propulsion, systems, and structures. Decisions that were made early in the planning of the aircraft project and the results of those decisions are briefly discussed. Each of the three flight phases corresponds to a particular aircraft configuration, and the research dictated the configuration to be flown. The first phase gathered data with the baseline F-18 configuration. The second phase was the thrust-vectoring phase. The third phase used a modified forebody with deployable nose strakes. Aircraft systems supporting these flights included extensive instrumentation systems, integrated research flight controls using flight control hardware and corresponding software, analog interface boxes to control forebody strakes, a thrust-vectoring system using external post-exit vanes around axisymmetric nozzles, a forebody vortex control system with strakes, and backup systems using battery-powered emergency systems and a spin recovery parachute.

Simulation Study of Flap Effects on a Commercial Transport Airplane in Upset Conditions

NASA Technical Reports Server (NTRS)

Cunningham, Kevin; Foster, John V.; Shah, Gautam H.; Stewart, Eric C.; Ventura, Robin N.; Rivers, Robert A.; Wilborn, James E.; Gato, William

2005-01-01

As part of NASA's Aviation Safety and Security Program, a simulation study of a twinjet transport airplane crew training simulation was conducted to address fidelity for upset or loss of control conditions and to study the effect of flap configuration in those regimes. Piloted and desktop simulations were used to compare the baseline crew training simulation model with an enhanced aerodynamic model that was developed for high-angle-of-attack conditions. These studies were conducted with various flap configurations and addressed the approach-to-stall, stall, and post-stall flight regimes. The enhanced simulation model showed that flap configuration had a significant effect on the character of departures that occurred during post-stall flight. Preliminary comparisons with flight test data indicate that the enhanced model is a significant improvement over the baseline. Some of the unrepresentative characteristics that are predicted by the baseline crew training simulation for flight in the post-stall regime have been identified. This paper presents preliminary results of this simulation study and discusses key issues regarding predicted flight dynamics characteristics during extreme upset and loss-of-control flight conditions with different flap configurations.

Lock-In Imaging System for Detecting Disturbances in Fluid

NASA Technical Reports Server (NTRS)

Park, Yeonjoon (Inventor); Choi, Sang Hyouk (Inventor); King, Glen C. (Inventor); Elliott, James R. (Inventor); Dimarcantonio, Albert L. (Inventor)

2014-01-01

A lock-in imaging system is configured for detecting a disturbance in air. The system includes an airplane, an interferometer, and a telescopic imaging camera. The airplane includes a fuselage and a pair of wings. The airplane is configured for flight in air. The interferometer is operatively disposed on the airplane and configured for producing an interference pattern by splitting a beam of light into two beams along two paths and recombining the two beams at a junction point in a front flight path of the airplane during flight. The telescopic imaging camera is configured for capturing an image of the beams at the junction point. The telescopic imaging camera is configured for detecting the disturbance in air in an optical path, based on an index of refraction of the image, as detected at the junction point.

Power Reactant Storage Assembly (PRSA) (Space Shuttle). PRSA hydrogen and oxygen DVT tank refurbishment

NASA Technical Reports Server (NTRS)

1993-01-01

The Power Reactant Storage Assembly (PRSA) liquid hydrogen Development Verification Test (H2 DVT) tank assembly (Beech Aircraft Corporation P/N 15548-0116-1, S/N 07399000SHT0001) and liquid oxygen (O2) DVT tank assembly (Beech Aircraft Corporation P/N 15548-0115-1, S/N 07399000SXT0001) were refurbished by Ball Electro-Optics and Cryogenics Division to provide NASA JSC, Propulsion and Power Division, the capability of performing engineering tests. The refurbishments incorporated the latest flight configuration hardware and avionics changes necessary to make the tanks function like flight articles. This final report summarizes these refurbishment activities. Also included are up-to-date records of the pressure time and cycle histories.

Saturn Apollo Program

NASA Image and Video Library

1968-03-03

The launch of the Apollo 9 (Saturn V launch vehicle, SA-504), with astronauts James A. McDivitt, David R. Scott, and Russell L. Schweickart, took place on March 3, 1968. The Apollo 9 spacecraft, in the lunar mission configuration, was tested in Earth orbit. The mission was designed to rehearse all the steps and reproduce all the events of the Apollo 11 mission with the exception of the lunar touchdown, stay, and liftoff. The command and service modules, and the lunar module were used in flight procedures identical to those that would later take similar vehicles to the Moon, and a landing. The flight mechanics, mission support systems, communications, and recording of data were tested in a final round of verification. Astronauts Scott and Schweickart conducted Extravehicular Activity during this mission.

Space Shuttle GN and C Development History and Evolution

NASA Technical Reports Server (NTRS)

Zimpfer, Douglas; Hattis, Phil; Ruppert, John; Gavert, Don

2011-01-01

Completion of the final Space Shuttle flight marks the end of a significant era in Human Spaceflight. Developed in the 1970 s, first launched in 1981, the Space Shuttle embodies many significant engineering achievements. One of these is the development and operation of the first extensive fly-by-wire human space transportation Guidance, Navigation and Control (GN&C) System. Development of the Space Shuttle GN&C represented first time inclusions of modern techniques for electronics, software, algorithms, systems and management in a complex system. Numerous technical design trades and lessons learned continue to drive current vehicle development. For example, the Space Shuttle GN&C system incorporated redundant systems, complex algorithms and flight software rigorously verified through integrated vehicle simulations and avionics integration testing techniques. Over the past thirty years, the Shuttle GN&C continued to go through a series of upgrades to improve safety, performance and to enable the complex flight operations required for assembly of the international space station. Upgrades to the GN&C ranged from the addition of nose wheel steering to modifications that extend capabilities to control of the large flexible configurations while being docked to the Space Station. This paper provides a history of the development and evolution of the Space Shuttle GN&C system. Emphasis is placed on key architecture decisions, design trades and the lessons learned for future complex space transportation system developments. Finally, some of the interesting flight operations experience is provided to inform future developers of flight experiences.

Veggie Processing

NASA Image and Video Library

2017-02-15

Charles Spern, at right, project manager on the Engineering Services Contract (ESC), and Glenn Washington, ESC quality assurance specialist, perform final inspections of the Veggie Series 1 plant experiment inside a laboratory in the Space Station Processing Facility at NASA's Kennedy Space Center in Florida. At far left is Dena Richmond, ESC configuration management. The Series 1 experiment is being readied for flight aboard Orbital ATK's Cygnus module on its seventh (OA-7) Commercial Resupply Services mission to the International Space Station. The Veggie system is on the space station.

Three Dimensional Aerodynamic Analysis of a High-Lift Transport Configuration

NASA Technical Reports Server (NTRS)

Dodbele, Simha S.

1993-01-01

Two computational methods, a surface panel method and an Euler method employing unstructured grid methodology, were used to analyze a subsonic transport aircraft in cruise and high-lift conditions. The computational results were compared with two separate sets of flight data obtained for the cruise and high-lift configurations. For the cruise configuration, the surface pressures obtained by the panel method and the Euler method agreed fairly well with results from flight test. However, for the high-lift configuration considerable differences were observed when the computational surface pressures were compared with the results from high-lift flight test. On the lower surface of all the elements with the exception of the slat, both the panel and Euler methods predicted pressures which were in good agreement with flight data. On the upper surface of all the elements the panel method predicted slightly higher suction compared to the Euler method. On the upper surface of the slat, pressure coefficients obtained by both the Euler and panel methods did not agree with the results of the flight tests. A sensitivity study of the upward deflection of the slat from the 40 deg. flap setting suggested that the differences in the slat deflection between the computational model and the flight configuration could be one of the sources of this discrepancy. The computation time for the implicit version of the Euler code was about 1/3 the time taken by the explicit version though the implicit code required 3 times the memory taken by the explicit version.

Spin reorientation of a nonsymmetric body with energy dissipation

NASA Technical Reports Server (NTRS)

Cenker, R. J.

1973-01-01

Stable rotating semi-rigid bodies were demonstrated analytically, and verified in flights such as Explorer 1 and ATS-5 satellites. The problem arises from the two potential orientations which the final spin vector can take after large angle reorientation from minor to major axis, i.e., along the positive or negative axis of the maximum inertia. Reorientation of a satellite initially spinning about the minor axis using an energy dissipation device may require that the final spin orientation be controlled. Examples of possible applications are the Apogee Motor Assembly with Paired Satellites (AMAPS) configuration, where proper orientation of the thruster is required; and reorientation of ATS-5, where the spin sensitive nature of the despin device (yo-yo mechanism) requires that the final spin vector point is a specified direction.

Free-Flight Investigation of Radio Controlled Models with Parawings

NASA Technical Reports Server (NTRS)

Hewes, Donald E.

1961-01-01

A free-flight investigation of two radio-controlled models with parawings, a glider configuration and an airplane (powered) configuration, was made to evaluate the performance, stability, and methods of controlling parawing vehicles. The flight tests showed that the models were stable and could be controlled either by shifting the center of gravity or by using conventional elevator and rudder control surfaces. Static wind-tunnel force-test data were also obtained.

Study of Alternate Space Shuttle Concepts. Volume 2, Part 2: Concept Analysis and Definition

NASA Technical Reports Server (NTRS)

1971-01-01

This is the final report of a Phase A Study of Alternate Space Shuttle Concepts by the Lockheed Missiles & Space Company (LMSC) for the National Aeronautics and Space Administration George C. Marshall Space Flight Center (MSFC). The eleven-month study, which began on 30 June 1970, is to examine the stage-and-one-half and other Space Shuttle configurations and to establish feasibility, performance, cost, and schedules for the selected concepts. This final report consists of four volumes as follows: Volume I - Executive Summary, Volume II - Concept Analysis and Definition, Volume III - Program Planning, and Volume IV - Data Cost Data. This document is Volume II, Concept Analysis and Definition.

Feasibility of Supersonic Aircraft Concepts for Low-Boom and Flight Trim Constraints

NASA Technical Reports Server (NTRS)

Li, Wu

2015-01-01

This paper documents a process for analyzing whether a particular supersonic aircraft configuration layout and a given cruise condition are feasible to achieve a trimmed low-boom design. This process was motivated by the need to know whether a particular configuration at a given cruise condition could be reshaped to satisfy both low-boom and flight trim constraints. Without such a process, much effort could be wasted on shaping a configuration layout at a cruise condition that could never satisfy both low-boom and flight trim constraints simultaneously. The process helps to exclude infeasible configuration layouts with minimum effort and allows a designer to develop trimmed low-boom concepts more effectively. A notional low-boom supersonic demonstrator concept is used to illustrate the analysis/design process.

A flight investigation of the wake turbulence alleviation resulting from a flap configuration change on a B-747 aircraft

NASA Technical Reports Server (NTRS)

Jacobsen, R. A.; Short, B. J.

1977-01-01

A flight test investigation was conducted to evaluate the effects of a flap configuration change on the vortex wake characteristics of a Boeing 747 (B-747) aircraft as measured by differences in upset response resulting from deliberate vortex encounters by a following Learjet aircraft and by direct measurement of the velocities in the wake. The flaps of the B-747 have a predominant effect on the wake. The normal landing flap configuration produces a strong vortex that is attenuated when the outboard flap segments are raised; however, extension of the landing gear at that point increases the vortex induced upsets. These effects are in general agreement with existing wind tunnel and flight data for the modified flap configuration.

Correlation of Predicted and Flight Derived Stability and Control Derivatives with Particular Application to Tailless Delta Wing Configurations

NASA Technical Reports Server (NTRS)

Weil, J.

1981-01-01

Flight derived longitudinal and lateral-directional stability and control derivatives were compared to wind-tunnel derived values. As a result of these comparisons, boundaries representing the uncertainties that could be expected from wind-tunnel predictions were established. These boundaries provide a useful guide for control system sensitivity studies prior to flight. The primary application for this data was the space shuttle, and as a result the configurations included in the study were those most applicable to the space shuttle. The configurations included conventional delta wing aircraft as well as the X-15 and lifting body vehicles.

14 CFR 23.251 - Vibration and buffeting.

Code of Federal Regulations, 2013 CFR

2013-01-01

... must be no buffeting in any normal flight condition, including configuration changes during cruise, severe enough to interfere with the satisfactory control of the airplane or cause excessive fatigue to... perceptible buffeting condition in the cruise configuration in straight flight at any speed up to VMO/MMO...

14 CFR 23.251 - Vibration and buffeting.

Code of Federal Regulations, 2014 CFR

2014-01-01

... must be no buffeting in any normal flight condition, including configuration changes during cruise, severe enough to interfere with the satisfactory control of the airplane or cause excessive fatigue to... perceptible buffeting condition in the cruise configuration in straight flight at any speed up to VMO/MMO...

Applied Meteorology Unit Quarterly Report. First Quarter FY-13

NASA Technical Reports Server (NTRS)

2013-01-01

The AMU team worked on five tasks for their customers: (1) Ms. Crawford continued work on the objective lightning forecast task for airports in east-central Florida. (2) Ms. Shafer continued work on the task for Vandenberg Air Force Base to create an automated tool that will help forecasters relate pressure gradients to peak wind values. (3) Dr. Huddleston began work to develop a lightning timing forecast tool for the Kennedy Space Center/Cape Canaveral Air Force Station area. (3) Dr. Bauman began work on a severe weather forecast tool focused on east-central Florida. (4) Dr. Watson completed testing high-resolution model configurations for Wallops Flight Facility and the Eastern Range, and wrote the final report containing the AMU's recommendations for model configurations at both ranges.

NASA rotor system research aircraft flight-test data report: Helicopter and compound configuration

NASA Technical Reports Server (NTRS)

Erickson, R. E.; Kufeld, R. M.; Cross, J. L.; Hodge, R. W.; Ericson, W. F.; Carter, R. D. G.

1984-01-01

The flight test activities of the Rotor System Research Aircraft (RSRA), NASA 740, from June 30, 1981 to August 5, 1982 are reported. Tests were conducted in both the helicopter and compound configurations. Compound tests reconfirmed the Sikorsky flight envelope except that main rotor blade bending loads reached endurance at a speed about 10 knots lower than previously. Wing incidence changes were made from 0 to 10 deg.

Effects of combining vertical and horizontal information into a primary flight display

NASA Technical Reports Server (NTRS)

Abbott, Terence S.; Nataupsky, Mark; Steinmetz, George G.

1987-01-01

A ground-based aircraft simulation study was conducted to determine the effects of combining vertical and horizontal flight information into a single display. Two display configurations were used in this study. The first configuration consisted of a Primary Flight Display (PFD) format and a Horizontal Situation Display (HSD) with the PFD displayed conventionally above the HSD. For the second display configuration, the HSD format was combined with the PFD format. Four subjects participated in this study. Data were collected on performance parameters, pilot-control inputs, auditory evoked response parameters (AEP), oculometer measurements (eye-scan), and heart rate. Subjective pilot opinion was gathered through questionnaire data and scorings for both the Subjective Workload Assessment Technique (SWAT) and the NASA Task Load Index (NASA-TLX). The results of this study showed that, from a performance and subjective standpoint, the combined configuration was better than the separate configuration. Additionally, both the eye-transition and eye-dwell times for the separate HSD were notably higher than expected, with a 46% increase in available visual time when going from double to single display configuration.

Stability and Control Estimation Flight Test Results for the SR-71 Aircraft With Externally Mounted Experiments

NASA Technical Reports Server (NTRS)

Moes, Timothy R.; Iliff, Kenneth

2002-01-01

A maximum-likelihood output-error parameter estimation technique is used to obtain stability and control derivatives for the NASA Dryden Flight Research Center SR-71A airplane and for configurations that include experiments externally mounted to the top of the fuselage. This research is being done as part of the envelope clearance for the new experiment configurations. Flight data are obtained at speeds ranging from Mach 0.4 to Mach 3.0, with an extensive amount of test points at approximately Mach 1.0. Pilot-input pitch and yaw-roll doublets are used to obtain the data. This report defines the parameter estimation technique used, presents stability and control derivative results, and compares the derivatives for the three configurations tested. The experimental configurations studied generally show acceptable stability, control, trim, and handling qualities throughout the Mach regimes tested. The reduction of directional stability for the experimental configurations is the most significant aerodynamic effect measured and identified as a design constraint for future experimental configurations. This report also shows the significant effects of aircraft flexibility on the stability and control derivatives.

Determination of longitudinal aerodynamic derivatives using flight data from an icing research aircraft

NASA Technical Reports Server (NTRS)

Ranaudo, R. J.; Batterson, J. G.; Reehorst, A. L.; Bond, T. H.; Omara, T. M.

1989-01-01

A flight test was performed with the NASA Lewis Research Center's DH-6 icing research aircraft. The purpose was to employ a flight test procedure and data analysis method, to determine the accuracy with which the effects of ice on aircraft stability and control could be measured. For simplicity, flight testing was restricted to the short period longitudinal mode. Two flights were flown in a clean (baseline) configuration, and two flights were flown with simulated horizontal tail ice. Forty-five repeat doublet maneuvers were performed in each of four test configurations, at a given trim speed, to determine the ensemble variation of the estimated stability and control derivatives. Additional maneuvers were also performed in each configuration, to determine the variation in the longitudinal derivative estimates over a wide range of trim speeds. Stability and control derivatives were estimated by a Modified Stepwise Regression (MSR) technique. A measure of the confidence in the derivative estimates was obtained by comparing the standard error for the ensemble of repeat maneuvers, to the average of the estimated standard errors predicted by the MSR program. A multiplicative relationship was determined between the ensemble standard error, and the averaged program standard errors. In addition, a 95 percent confidence interval analysis was performed for the elevator effectiveness estimates, C sub m sub delta e. This analysis identified the speed range where changes in C sub m sub delta e could be attributed to icing effects. The magnitude of icing effects on the derivative estimates were strongly dependent on flight speed and aircraft wing flap configuration. With wing flaps up, the estimated derivatives were degraded most at lower speeds corresponding to that configuration. With wing flaps extended to 10 degrees, the estimated derivatives were degraded most at the higher corresponding speeds. The effects of icing on the changes in longitudinal stability and control derivatives were adequately determined by the flight test procedure and the MSR analysis method discussed herein.

Trans Atlantic Infrasound Payload (TAIP) Operation Plan.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bowman, Daniel; Lees, Jonathan M.

The Carolina Infrasound package, added as a piggyback to the 2016 ULDB ight, recorded unique acoustic signals such as the ocean microbarom and a large meteor. These data both yielded unique insights into the acoustic energy transfer from the lower to the upper atmosphere as well as highlighted the vast array of signals whose origins remain unknown. Now, the opportunity to y a payload across the north Atlantic offers an opportunity to sample one of the most active ocean microbarom sources on Earth. Improvements in payload capabilities should result in characterization of the higher frequency range of the stratospheric infrasoundmore » spectrum as well. Finally, numerous large mining and munitions disposal explosions in the region may provide \\ground truth" events for assessing the detection capability of infrasound microphones in the stratosphere. The flight will include three different types of infrasound sensors. One type is a pair of polarity reversed InfraBSU microphones (standard for high altitude flights since 2016), another is a highly sensitive Chaparral 60 modified for a very low corner period, and the final sensor is a lightweight, low power Gem infrasound package. By evaluating these configurations against each other on the same flight, we will be able to optimize future campaigns with different sensitivity and mass constraints.« less

Aerodynamic Characteristics and Flying Qualities of a Tailless Triangular-wing Airplane Configuration as Obtained from Flights of Rocket-propelled Models at Transonic and Supersonic Speeds

NASA Technical Reports Server (NTRS)

Mitcham, Grady L; Stevens, Joseph E; Norris, Harry P

1956-01-01

A flight investigation of rocket-powered models of a tailless triangular-wing airplane configuration was made through the transonic and low supersonic speed range at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. An analysis of the aerodynamic coefficients, stability derivatives, and flying qualities based on the results obtained from the successful flight tests of three models is presented.

Beam optical design of in-flight fragment separator for high-power heavy ion beam

NASA Astrophysics Data System (ADS)

Yun, C. C.; Kim, Mi-Jung; Kim, D. G.; Song, J. S.; Kim, Myeong-Jin; Kim, J. W.; Kim, J. R.; Wan, W.

2013-12-01

An in-flight fragment separator has been designed for the rare isotope science project (RISP) in Korea. A beam used for the design is 238U in the energy of 200 MeV/u with the maximum beam power of 400 kW. The use of high-power beam requires careful removal of the primary beam by pre-separator, for which its configuration was revised to employ four dipole magnets instead of two. Different configurations of the separator have been tested in search of optimal design in non-linear optics, which was complicated by the space needed for the target, beam dump and radiation shielding. Non-linear optical calculations have been carried out using GICOSY and COSY Infinity including the fringe fields of large-aperture quadrupole magnets. Correction of non-linear terms is made with multipole coils located inside the superconducting quadrupole magnets and by external multipole magnets. Beam simulations using LISE++ and MOCADI have been performed to consider the effects of multiple charge states of the primary and isotope beams produced at the target. Layout of the separator is being finalized, and detailed optics simulation will continue to refine its design.

Abort Flight Test Project Overview

NASA Technical Reports Server (NTRS)

Sitz, Joel

2007-01-01

A general overview of the Orion abort flight test is presented. The contents include: 1) Abort Flight Test Project Overview; 2) DFRC Exploration Mission Directorate; 3) Abort Flight Test; 4) Flight Test Configurations; 5) Flight Test Vehicle Engineering Office; 6) DFRC FTA Scope; 7) Flight Test Operations; 8) DFRC Ops Support; 9) Launch Facilities; and 10) Scope of Launch Abort Flight Test

KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Jacobs, P.; Flechner, S.; Sims, R.

1982-01-01

A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

FLITECAM: delivery and performance on SOFIA

NASA Astrophysics Data System (ADS)

Logsdon, Sarah E.; McLean, Ian S.; Becklin, E. E.; Hamilton, Ryan T.; Vacca, William D.; Waddell, Patrick

2016-08-01

We present a performance report for FLITECAM, a 1-5 μm imager and spectrograph, upon its acceptance and delivery to SOFIA (Stratospheric Observatory for Infrared Astronomy). FLITECAM has two observing configurations: solo configuration and "FLIPO" configuration, which is the co-mounting of FLITECAM with the optical instrument HIPO (PI E. Dunham, Lowell Observatory). FLITECAM was commissioned in the FLIPO configuration in 2014 and flew in the solo configuration for the first time in Fall 2015, shortly after its official delivery to SOFIA. Here we quantify FLITECAM's imaging and spectral performance in both configurations and discuss the science capabilities of each configuration, with examples from in-flight commissioning and early science data. The solo configuration (which comprises fewer warm optics) has better sensitivity at longer wavelengths. We also discuss the causes of excess background detected in the in-flight FLITECAM images at low elevations and describe the current plan to mitigate the largest contributor to this excess background.

HIFiRE-5 Flight Vehicle Design

NASA Technical Reports Server (NTRS)

Kimmel, Roger L.; Adamczak, David; Berger, Karen; Choudhari, Meelan

2010-01-01

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratories (AFRL) and Australian Defence Science and Technology Organization (DSTO). HIFiRE flight 5 is devoted to measuring transition on a three-dimensional body. This paper summarizes payload configuration, trajectory, vehicle stability limits and roughness tolerances. Results show that the proposed configuration is suitable for testing transition on a three-dimensional body. Transition is predicted to occur within the test window, and a design has been developed that will allow the vehicle to be manufactured within prescribed roughness tolerances

Aft-End Flow of a Large-Scale Lifting Body During Free-Flight Tests

NASA Technical Reports Server (NTRS)

Banks, Daniel W.; Fisher, David F.

2006-01-01

Free-flight tests of a large-scale lifting-body configuration, the X-38 aircraft, were conducted using tufts to characterize the flow on the aft end, specifically in the inboard region of the vertical fins. Pressure data was collected on the fins and base. Flow direction and movement were correlated with surface pressure and flight condition. The X-38 was conceived to be a rescue vehicle for the International Space Station. The vehicle shape was derived from the U.S. Air Force X-24 lifting body. Free-flight tests of the X-38 configuration were conducted at the NASA Dryden Flight Research Center at Edwards Air Force Base, California from 1997 to 2001.

Space Shuttle solid rocket motor /SRM/ development and qualification

NASA Technical Reports Server (NTRS)

Lund, R. K.; Brinton, B. C.

1980-01-01

The configuration of the reusable Space Shuttle solid rocket motors is described. In addition, their design evolution is reviewed, noting that the requirement that certain components be recovered, refurbished, and used on as many as 20 flights dictated a conservative design approach, the validity of which has been proven by successful testing of all development and qualification motors. Aspects discussed include ballistics, the motor case, nozzle, nozzle materials, and the ignition system. Finally, summary results of the first two of three qualification motor firings designated QM-1 and QM-2 are presented.

Experimental investigation of a quad-rotor biplane micro air vehicle

NASA Astrophysics Data System (ADS)

Bogdanowicz, Christopher Michael

Micro air vehicles are expected to perform demanding missions requiring efficient operation in both hover and forward flight. This thesis discusses the development of a hybrid air vehicle which seamlessly combines both flight capabilities: hover and high-speed forward flight. It is the quad-rotor biplane, which weighs 240 grams and consists of four propellers with wings arranged in a biplane configuration. The performance of the vehicle system was investigated in conditions representative of flight through a series of wind tunnel experiments. These studies provided an understanding of propeller-wing interaction effects and system trim analysis. This showed that the maximum speed of 11 m/s and a cruise speed of 4 m/s were achievable and that the cruise power is approximately one-third of the hover power. Free flight testing of the vehicle successfully highlighted its ability to achieve equilibrium transition flight. Key design parameters were experimentally investigated to understand their effect on overall performance. It was found that a trade-off between efficiency and compactness affects the final choice of the design. Design improvements have allowed for decreases in vehicle weight and ground footprint, while increasing structural soundness. Numerous vehicle designs, models, and flight tests have proven system scalability as well as versatility, including an upscaled model to be utilized in an extensive commercial package delivery system. Overall, the quad-rotor biplane is proven to be an efficient and effective multi-role vehicle.

Analysis of the Flight Motions of a Small Deployable Glider Configuration

NASA Technical Reports Server (NTRS)

Coe, Paul L., Jr.

1975-01-01

An investigation was conducted at the request of the U.S. Air Force Avionics Laboratory to analyze the flight characteristics of a small uncontrolled glider with folding wings. The study consisted of wind-tunnel tests of an actual glider and a theoretical analysis of the performance, stability, and trimmability of the configuration.

14 CFR 417.125 - Launch of an unguided suborbital launch vehicle.

Code of Federal Regulations, 2010 CFR

2010-01-01

... elevation angle setting that ensures the rocket will not fly uprange. A launch operator must set the... throughout each stage of powered flight. A caliber, for a rocket configuration, is defined as the distance... rocket configuration. (f) Tracking. A launch operator must track the flight of an unguided suborbital...

14 CFR 417.125 - Launch of an unguided suborbital launch vehicle.

Code of Federal Regulations, 2013 CFR

2013-01-01

... elevation angle setting that ensures the rocket will not fly uprange. A launch operator must set the... throughout each stage of powered flight. A caliber, for a rocket configuration, is defined as the distance... rocket configuration. (f) Tracking. A launch operator must track the flight of an unguided suborbital...

14 CFR 417.125 - Launch of an unguided suborbital launch vehicle.

Code of Federal Regulations, 2012 CFR

2012-01-01

... elevation angle setting that ensures the rocket will not fly uprange. A launch operator must set the... throughout each stage of powered flight. A caliber, for a rocket configuration, is defined as the distance... rocket configuration. (f) Tracking. A launch operator must track the flight of an unguided suborbital...

14 CFR 417.125 - Launch of an unguided suborbital launch vehicle.

Code of Federal Regulations, 2011 CFR

2011-01-01

... elevation angle setting that ensures the rocket will not fly uprange. A launch operator must set the... throughout each stage of powered flight. A caliber, for a rocket configuration, is defined as the distance... rocket configuration. (f) Tracking. A launch operator must track the flight of an unguided suborbital...

14 CFR 417.125 - Launch of an unguided suborbital launch vehicle.

Code of Federal Regulations, 2014 CFR

2014-01-01

... elevation angle setting that ensures the rocket will not fly uprange. A launch operator must set the... throughout each stage of powered flight. A caliber, for a rocket configuration, is defined as the distance... rocket configuration. (f) Tracking. A launch operator must track the flight of an unguided suborbital...

An in-flight investigation of a twin fuselage configuration in approach and landing

NASA Technical Reports Server (NTRS)

Weingarten, N. C.

1984-01-01

An in-flight investigation of the flying qualities of a twin fuselage aircraft design in the approach and landing flight phase was carried out in the USAF/AFWAL Total In-Flight Simulator (TIFS). The objective was to determine the effects of actual motion and visual cues on the pilot when he was offset from the centerline of the aircraft. The experiment variables were lateral pilot offset position (0, 30 and 50 feet) and effective roll mode time constant (.6, 1.2, 2.4 seconds). The evaluation included the final approach, flare and touchdown. Lateral runway offsets and 15 knot crosswinds were used to increase the pilot's workload and force him to make large lateral corrections in the final portion of the approach. Results indicated that large normal accelerations rather than just vertical displacements in rolling maneuvers had the most significant degrading effect on pilot ratings. The normal accelerations are a result of large lateral offset and fast roll mode time constant and caused the pilot to make unnecessary pitch inputs and get into a coupled pitch/roll oscillation while he was making line up and crosswind corrections. A potential criteria for lateral pilot offset position effects is proposed. When the ratio of incremented normal aceleration at the pilot station to the steady state roll rate for a step input reaches .01 to .02 g/deg/sec a deterioration of pilot rating and flying qualities level can be expected.

Cobra communications switch integration program

NASA Technical Reports Server (NTRS)

Shively, Robert J.; Haworth, Loran A.; Szoboszlay, Zoltan; Murray, F. Gerald

1989-01-01

The paper describes a design modification to reduce the visual and manual workload associated with the radio selection and communications tasks in the U.S. Army AH-1 Cobra helicopter. The modification involves the integration of the radio selection and microphone actuating tasks into a single operation controlled by the transmit-intercom switch. Ground-based and flight tests were conducted to evaluate the modified configuration during twelve flight tasks. The results show that the proposed configuration performs twice as fast as the original configuration.

Conceptual Design of a Tiltrotor Transport Flight Deck

NASA Technical Reports Server (NTRS)

Decker, William A.; Dugan, Daniel C.; Simmons, Rickey C.; Tucker, George E.; Aiken, Edwin W. (Technical Monitor)

1995-01-01

A tiltrotor transport has considerable potential as a regional transport, increasing the air transportation system capacity by off-loading conventional runways. Such an aircraft will have a flight deck suited to its air transportation task and adapted to unique urban vertiport operating requirements. Such operations are likely to involve steep, slow instrument approaches for vertical and extremely short rolling take-offs and landings. While much of a tiltrotor transport's operations will be in common with commercial fixed-wing operations, terminal area operations will impose alternative flight deck design solutions. Control systems, displays and guidance, and control inceptors must be tailored to both routine and emergency vertical flight operations. This paper will survey recent experience with flight deck design elements suitable to a tiltrotor transport and will propose a conceptual cockpit design for such an aircraft. A series of piloted simulations using the NASA Ames Vertical Motion Simulator have investigated cockpit design elements and operating requirements for tiltrotor transports operating into urban vertiports. These experiments have identified the need for a flight director or equivalent display guidance for steep final approaches. A flight path vector display format has proven successful for guiding tiltrotor transport terminal area operations. Experience with a Head-Up Display points to the need for a bottom-mounted display device to maximize its utility on steep final approach paths. Configuration control (flap setting and nacelle angle) requires appropriate augmentation and tailoring for civil transport operations, flown to an airline transport pilot instrument flight rules (ATP-IFR) standard. The simulation experiments also identified one thrust control lever geometry as inappropriate to the task and found at least acceptable results with the vertical thrust control lever of the XV-15. In addition to the thrust controller, the attitude control of a tiltrotor transport may be effected through an inceptor other than the current center sticks in the XV-15 and V-22. Simulation and flight investigations of side-stick control inceptors for rotorcraft, augmented by a 1985 flight test of a side-stick controller in the XV-15 suggest the potential of such a device in a transport cockpit.

Design of a flight director/configuration management system for piloted STOL approaches

NASA Technical Reports Server (NTRS)

Hoh, R. H.; Klein, R. H.; Johnson, W. A.

1973-01-01

The design and characteristics of a flight director for V/STOL aircraft are discussed. A configuration management system for piloted STOL approaches is described. The individual components of the overall system designed to reduce pilot workload to an acceptable level during curved, decelerating, and descending STOL approaches are defined. The application of the system to augmentor wing aircraft is analyzed. System performance checks and piloted evaluations were conducted on a flight simulator and the results are summarized.

NASA rotor systems research aircraft: Fixed-wing configuration flight-test results

NASA Technical Reports Server (NTRS)

Erickson, R. E.; Cross, J. L.; Kufeld, R. M.; Acree, C. W.; Nguyen, D.; Hodge, R. W.

1986-01-01

The fixed-wing, airplane configuration flight-test results of the Rotor System Research Aircraft (RSRA), NASA 740, at Ames/Dryden Flight Research Center are documented. Fourteen taxi and flight tests were performed from December 1983 to October 1984. This was the first time the RSRA was flown with the main rotor removed; the tail rotor was installed. These tests confirmed that the RSRA is operable as a fixed-wing aircraft. Data were obtained for various takeoff and landing distances, control sensitivity, trim and dynamics stability characteristics, performance rotor-hub drag, and acoustics signature. Stability data were obtained with the rotor hub both installed and removed. The speed envelope was developed to 261 knots true airspeed (KTAS), 226 knots calibrated airspeed (KCAS) at 10,000 ft density altitude. The airplane was configured at 5 deg. wing incidence with 5 deg. wing flaps as a normal configuration. Level-flight data were acquired at 167 KCAS for wing incidence from 0 to 10 deg. Step inputs and doublet inputs of various magnitudes were utilized to acquire dynamic stability and control sensitivity data. Sine-wave inputs of constantly increasing frequency were used to generate parameter identification data. The maximum load factor attained was 2.34 g at 206 KCAS.

Design, Development and Testing of Airplanes for Mars Exploration

NASA Technical Reports Server (NTRS)

Hall, David W.

2004-01-01

The opportunity for a piggyback mission to Mars aboard an Ariane 5 rocket in the early spring of 1999 set off feverish design activity at several NASA centers. This report describes the contract work done by faculty, students, and consultants at the California Polytechnic State University in San Luis Obispo California (Cal poly/SLO) to support the NASA/Ames design, construction and test efforts to develop a simple and robust Mars Flyer configuration capable of performing a practical science mission on Mars. The first sections will address the conceptual design of a workable Mars Flyer configuration which started in the spring and summer of 1999. The following sections will focus on construction and flight test of two full-scale vehicles. The final section will reflect on the overall effort and make recommendations for future work.

High Detent Torque Rotary Actuator Development

NASA Astrophysics Data System (ADS)

Santos, I.; Sainz, I.; Allegranza, C.

2015-09-01

In the frame of an ESA ARTES 5 Contract, SENER has performed the design, manufacturing and testing at component and mechanism levels of a High Detent Torque Rotary Actuator (DTA in short), i.e. with high capability to hold a payload when unpowered.Two configurations were developed to allow the use on specific application flight opportunity; both are identical in terms of architecture, lubrication, structural and thermal design. The exception is the angular position sensor type: the DTA 100 with contactless sensors and the DTA 120 with potentiometers.The DTA is a fully european technology. This paper provides a synthesis of the obtained parameters in front of the requirements, the evolution from the initial concept to the final configuration and the results of the extensive test campaign (DTA 120). Lessons learned and the readiness for use at upper level are also highlighted.

Space Shuttle Main Engine - The Relentless Pursuit of Improvement

NASA Technical Reports Server (NTRS)

VanHooser, Katherine P.; Bradley, Douglas P.

2011-01-01

The Space Shuttle Main Engine (SSME) is the only reusable large liquid rocket engine ever developed. The specific impulse delivered by the staged combustion cycle, substantially higher than previous rocket engines, minimized volume and weight for the integrated vehicle. The dual pre-burner configuration permitted precise mixture ratio and thrust control while the fully redundant controller and avionics provided a very high degree of system reliability and health diagnosis. The main engine controller design was the first rocket engine application to incorporate digital processing. The engine was required to operate at a high chamber pressure to minimize engine volume and weight. Power level throttling was required to minimize structural loads on the vehicle early in flight and acceleration levels on the crew late in ascent. Fatigue capability, strength, ease of assembly and disassembly, inspectability, and materials compatibility were all major considerations in achieving a fully reusable design. During the multi-decade program the design evolved substantially using a series of block upgrades. A number of materials and manufacturing challenges were encountered throughout SSME s history. Significant development was required for the final configuration of the high pressure turbopumps. Fracture control was implemented to assess life limits of critical materials and components. Survival in the hydrogen environment required assessment of hydrogen embrittlement. Instrumentation systems were a challenge due to the harsh thermal and dynamic environments within the engine. Extensive inspection procedures were developed to assess the engine components between flights. The Space Shuttle Main Engine achieved a remarkable flight performance record. All flights were successful with only one mission requiring an ascent abort condition, which still resulted in an acceptable orbit and mission. This was achieved in large part via extensive ground testing to fully characterize performance and to establish acceptable life limits. During the program over a million seconds of accumulated test and flight time was achieved. Post flight inspection and assessment was a key part of assuring proper performance of the flight hardware. By the end of the program the predicted reliability had improved by a factor of four. These unique challenges, evolution of the design, and the resulting reliability will be discussed in this paper.

Flight Telerobotic Servicer prototype simulator

NASA Astrophysics Data System (ADS)

Schein, Rob; Krauze, Linda; Hartley, Craig; Dickenson, Alan; Lavecchia, Tom; Working, Bob

A prototype simulator for the Flight Telerobotic Servicer (FTS) system is described for use in the design development of the FTS, emphasizing the hand controller and user interface. The simulator utilizes a graphics workstation based on rapid prototyping tools for systems analyses of the use of the user interface and the hand controller. Kinematic modeling, manipulator-control algorithms, and communications programs are contained in the software for the simulator. The hardwired FTS panels and operator interface for use on the STS Orbiter are represented graphically, and the simulated controls function as the final FTS system configuration does. The robotic arm moves based on the user hand-controller interface, and the joint angles and other data are given on the prototype of the user interface. This graphics simulation tool provides the means for familiarizing crewmembers with the FTS system operation, displays, and controls.

Time-Spectral Rotorcraft Simulations on Overset Grids

NASA Technical Reports Server (NTRS)

Leffell, Joshua I.; Murman, Scott M.; Pulliam, Thomas H.

2014-01-01

The Time-Spectral method is derived as a Fourier collocation scheme and applied to NASA's overset Reynolds-averaged Navier-Stokes (RANS) solver OVERFLOW. The paper outlines the Time-Spectral OVERFLOWimplementation. Successful low-speed laminar plunging NACA 0012 airfoil simulations demonstrate the capability of the Time-Spectral method to resolve the highly-vortical wakes typical of more expensive three-dimensional rotorcraft configurations. Dealiasing, in the form of spectral vanishing viscosity (SVV), facilitates the convergence of Time-Spectral calculations of high-frequency flows. Finally, simulations of the isolated V-22 Osprey tiltrotor for both hover and forward (edgewise) flight validate the three-dimensional Time-Spectral OVERFLOW implementation. The Time-Spectral hover simulation matches the time-accurate calculation using a single harmonic. Significantly more temporal modes and SVV are required to accurately compute the forward flight case because of its more active, high-frequency wake.

Flight experiments to improve terminal area operations

NASA Technical Reports Server (NTRS)

Salmirs, S.; Morello, S. A.

1978-01-01

A brief description is given of the objectives and activities of the terminal configured vehicle (TCV) program and of some of the airborne facilities. A short analysis of some particular problems in CTOL operations in the terminal area is also presented to show how the program's technical objectives are related to the defined problems. The test aircraft was flown both manually and automatically with manual monitoring over paths including 130 deg intercepts and 2.0 km and 0.8 km finals. Some statistical data are presented from these and other flight profiles designed to address specific terminal area problems. An overview is presented of research studies receiving emphasis in the next biennium and their application to the terminal area. A description of work undertaken to study the addition of adjacent traffic information to present map displays is also given.

Flight experiments to improve terminal area operations

NASA Technical Reports Server (NTRS)

Salmirs, S.; Morello, S. A.

1978-01-01

A brief description is given of the objectives and activities of the terminal configured vehicle (TCV) program and of some of the airborne facilities. A short analysis of some particular problems of CTOL operations in the terminal area is also presented to show how the program's technical objectives are related to the defined problems. The test aircraft was flown both manually and automatically with manual monitoring over paths including 130 deg intercepts and 2.0 km (1.1. n. mi.) and 0.8 km (0.44 n. mi.) finals. Some statistical data are presented from these and other flight profiles designed to address specific terminal in the next biennium and their application to the terminal area. A description of work being undertaken to study the addition of adjacent traffic information to present map displays is also given.

B-52 control configured vehicles: Flight test results

NASA Technical Reports Server (NTRS)

Arnold, J. I.; Murphy, F. B.

1976-01-01

Recently completed B-52 Control Configured Vehicles (CCV) flight testing is summarized, and results are compared to analytical predictions. Results are presented for five CCV system concepts: ride control, maneuver load control, flutter mode control, augmented stability, and fatigue reduction. Test results confirm analytical predictions and show that CCV system concepts achieve performance goals when operated individually or collectively.

Integrated Flight and Propulsion Controls for Advanced Aircraft Configurations

NASA Technical Reports Server (NTRS)

Merrill, Walter; Garg, Sanjay

1995-01-01

The research vision of the NASA Lewis Research Center in the area of integrated flight and propulsion controls technologies is described. In particular the Integrated Method for Propulsion and Airframe Controls developed at the Lewis Research Center is described including its application to an advanced aircraft configuration. Additionally, future research directions in integrated controls are described.

Integrated Flight and Propulsion Controls for Advanced Aircraft Configurations

NASA Technical Reports Server (NTRS)

Merrill, Walter; Garg, Sanjay

1996-01-01

The research vision of the NASA Lewis Research Center in the area of integrated flight and propulsion controls technologies is described. In particular, the integrated method for propulsion and airframe controls developed at the Lewis Research Center is described including its application to an advanced aircraft configuration. Additionally, future research directions in integrated controls are described.

Flight service evaluation of PRD-49/epoxy composite panels in wide-bodied commercial transport aircraft

NASA Technical Reports Server (NTRS)

Wooley, J. H.

1974-01-01

Fairing panels were fabricated to evaluate the fabrication characteristics and flight service performance of PRD-49 (Kevlar-49) a composite reinforcing material and to compare it with the fiberglass which is currently in use. Panel configurations were selected to evaluate the PRD-49 with two resin matrix materials in sandwich and solid laminate construction. Left and right hand versions of these configurations were installed on L-1011's which will accumulate approximately 3000 flight hours per year per aircraft. The direct substitution of PRD-49 for fiberglass produced a twenty-six percent weight reduction on the panel configurations. Examination of these panels revealed that there was no visible difference between the PRD-49 and adjacent fiberglass panels.

Intelligent redundant actuation system requirements and preliminary system design

NASA Technical Reports Server (NTRS)

Defeo, P.; Geiger, L. J.; Harris, J.

1985-01-01

Several redundant actuation system configurations were designed and demonstrated to satisfy the stringent operational requirements of advanced flight control systems. However, this has been accomplished largely through brute force hardware redundancy, resulting in significantly increased computational requirements on the flight control computers which perform the failure analysis and reconfiguration management. Modern technology now provides powerful, low-cost microprocessors which are effective in performing failure isolation and configuration management at the local actuator level. One such concept, called an Intelligent Redundant Actuation System (IRAS), significantly reduces the flight control computer requirements and performs the local tasks more comprehensively than previously feasible. The requirements and preliminary design of an experimental laboratory system capable of demonstrating the concept and sufficiently flexible to explore a variety of configurations are discussed.

Preliminary system design study for a digital fly-by-wire flight control system for an F-8C aircraft

NASA Technical Reports Server (NTRS)

Seacord, C. L.; Vaughn, D. K.

1976-01-01

The design of a fly-by-wire control system having a mission failure probability of less than one millionth failures per flight hour is examined. Emphasis was placed on developing actuator configurations that would improve the system performance, and consideration of the practical aspects of sensor/computer and computer/actuator interface implementation. Five basic configurations were defined as appropriate candidates for the F-8C research aircraft. Options on the basic configurations were included to cover variations in flight sensors, redundancy levels, data transmission techniques, processor input/output methods, and servo actuator arrangements. The study results can be applied to fly by wire systems for transport aircraft in general and the space shuttle.

Configuring the Orion Guidance, Navigation, and Control Flight Software for Automated Sequencing

NASA Technical Reports Server (NTRS)

Odegard, Ryan G.; Siliwinski, Tomasz K.; King, Ellis T.; Hart, Jeremy J.

2010-01-01

The Orion Crew Exploration Vehicle is being designed with greater automation capabilities than any other crewed spacecraft in NASA s history. The Guidance, Navigation, and Control (GN&C) flight software architecture is designed to provide a flexible and evolvable framework that accommodates increasing levels of automation over time. Within the GN&C flight software, a data-driven approach is used to configure software. This approach allows data reconfiguration and updates to automated sequences without requiring recompilation of the software. Because of the great dependency of the automation and the flight software on the configuration data, the data management is a vital component of the processes for software certification, mission design, and flight operations. To enable the automated sequencing and data configuration of the GN&C subsystem on Orion, a desktop database configuration tool has been developed. The database tool allows the specification of the GN&C activity sequences, the automated transitions in the software, and the corresponding parameter reconfigurations. These aspects of the GN&C automation on Orion are all coordinated via data management, and the database tool provides the ability to test the automation capabilities during the development of the GN&C software. In addition to providing the infrastructure to manage the GN&C automation, the database tool has been designed with capabilities to import and export artifacts for simulation analysis and documentation purposes. Furthermore, the database configuration tool, currently used to manage simulation data, is envisioned to evolve into a mission planning tool for generating and testing GN&C software sequences and configurations. A key enabler of the GN&C automation design, the database tool allows both the creation and maintenance of the data artifacts, as well as serving the critical role of helping to manage, visualize, and understand the data-driven parameters both during software development and throughout the life of the Orion project.

A Look at Handling Qualities of Canard Configurations

NASA Technical Reports Server (NTRS)

Anderson, Seth B.

1986-01-01

The first human-powered flight was achieved by a canard-configured aircraft (Wright Brothers). Although other canard concepts were flown with varying degrees of success over the years, the tail-aft configuration has dominated the aircraft market for both military and civil use. Reviewed are the development of several canard aircraft with emphasis on stability and control, handling qualities, and operating problems. The results show that early canard concepts suffered adversely in flight behavior because of a lack of understanding of the sensitivities of these concepts to basic stability and control principles. Modern canard designs have been made competitive with tail-aft configurations by using appropriate handling qualities design criteria.

Design, fabrication & performance analysis of an unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Khan, M. I.; Salam, M. A.; Afsar, M. R.; Huda, M. N.; Mahmud, T.

2016-07-01

An Unmanned Aerial Vehicle was designed, analyzed and fabricated to meet design requirements and perform the entire mission for an international aircraft design competition. The goal was to have a balanced design possessing, good demonstrated flight handling qualities, practical and affordable manufacturing requirements while providing a high vehicle performance. The UAV had to complete total three missions named ferry flight (1st mission), maximum load mission (2nd mission) and emergency medical mission (3rd mission). The requirement of ferry flight mission was to fly as many as laps as possible within 4 minutes. The maximum load mission consists of flying 3 laps while carrying two wooden blocks which simulate cargo. The requirement of emergency medical mission was complete 3 laps as soon as possible while carrying two attendances and two patients. A careful analysis revealed lowest rated aircraft cost (RAC) as the primary design objective. So, the challenge was to build an aircraft with minimum RAC that can fly fast, fly with maximum payload, and fly fast with all the possible configurations. The aircraft design was reached by first generating numerous design concepts capable of completing the mission requirements. In conceptual design phase, Figure of Merit (FOM) analysis was carried out to select initial aircraft configuration, propulsion, empennage and landing gear. After completion of the conceptual design, preliminary design was carried out. The preliminary design iterations had a low wing loading, high lift coefficient, and a high thrust to weight ratio. To make the aircraft capable of Rough Field Taxi; springs were added in the landing gears for absorbing shock. An airfoil shaped fuselage was designed to allowed sufficient space for payload and generate less drag to make the aircraft fly fast. The final design was a high wing monoplane with conventional tail, single tractor propulsion system and a tail dragger landing gear. Payload was stored in undercarriage box for maximum load mission and emergency medical mission. The aircraft structure, weights 5.6 lb., constructed by balsa wood, depron and covering film was the only feasible match for the given requirements set by the competition organizers. The defined final aircraft was capable of: Completing 3 laps within 4 minutes at the first mission; flying 3 laps with 4 internal payloads at the second mission; flying 3 laps with all possible payload configurations at the third mission.

Linear Aerospike SR-71 Experiment (LASRE) refueling during first flight

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 refuels with an Edwards Air Force Base KC-135 during the first flight of the NASA/Rocketdyne/ Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE). The flight took place Oct. 31 at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

Spherical primary optical telescope (SPOT) segments

NASA Astrophysics Data System (ADS)

Hall, Christopher; Hagopian, John; DeMarco, Michael

2012-09-01

The spherical primary optical telescope (SPOT) project is an internal research and development program at NASA Goddard Space Flight Center. The goals of the program are to develop a robust and cost effective way to manufacture spherical mirror segments and demonstrate a new wavefront sensing approach for continuous phasing across the segmented primary. This paper focuses on the fabrication of the mirror segments. Significant cost savings were achieved through the design, since it allowed the mirror segments to be cast rather than machined from a glass blank. Casting was followed by conventional figuring at Goddard Space Flight Center. After polishing, the mirror segments were mounted to their composite assemblies. QED Technologies used magnetorheological finishing (MRF®) for the final figuring. The MRF process polished the mirrors while they were mounted to their composite assemblies. Each assembly included several magnetic invar plugs that extended to within an inch of the face of the mirror. As part of this project, the interaction between the MRF magnetic field and invar plugs was evaluated. By properly selecting the polishing conditions, MRF was able to significantly improve the figure of the mounted segments. The final MRF figuring demonstrates that mirrors, in the mounted configuration, can be polished and tested to specification. There are significant process capability advantes due to polishing and testing the optics in their final, end-use assembled state.

Ground Data System Analysis Tools to Track Flight System State Parameters for the Mars Science Laboratory (MSL) and Beyond

NASA Technical Reports Server (NTRS)

Allard, Dan; Deforrest, Lloyd

2014-01-01

Flight software parameters enable space mission operators fine-tuned control over flight system configurations, enabling rapid and dynamic changes to ongoing science activities in a much more flexible manner than can be accomplished with (otherwise broadly used) configuration file based approaches. The Mars Science Laboratory (MSL), Curiosity, makes extensive use of parameters to support complex, daily activities via commanded changes to said parameters in memory. However, as the loss of Mars Global Surveyor (MGS) in 2006 demonstrated, flight system management by parameters brings with it risks, including the possibility of losing track of the flight system configuration and the threat of invalid command executions. To mitigate this risk a growing number of missions have funded efforts to implement parameter tracking parameter state software tools and services including MSL and the Soil Moisture Active Passive (SMAP) mission. This paper will discuss the engineering challenges and resulting software architecture of MSL's onboard parameter state tracking software and discuss the road forward to make parameter management tools suitable for use on multiple missions.

Peak Seeking Control for Reduced Fuel Consumption with Preliminary Flight Test Results

NASA Technical Reports Server (NTRS)

Brown, Nelson

2012-01-01

The Environmentally Responsible Aviation project seeks to accomplish the simultaneous reduction of fuel burn, noise, and emissions. A project at NASA Dryden Flight Research Center is contributing to ERAs goals by exploring the practical application of real-time trim configuration optimization for enhanced performance and reduced fuel consumption. This peak-seeking control approach is based on Newton-Raphson algorithm using a time-varying Kalman filter to estimate the gradient of the performance function. In real-time operation, deflection of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of a modified F-18 are directly optimized, and the horizontal stabilators and angle of attack are indirectly optimized. Preliminary results from three research flights are presented herein. The optimization system found a trim configuration that required approximately 3.5% less fuel flow than the baseline trim at the given flight condition. The algorithm consistently rediscovered the solution from several initial conditions. These preliminary results show the algorithm has good performance and is expected to show similar results at other flight conditions and aircraft configurations.

A new low-turbulence wind tunnel for animal and small vehicle flight experiments

NASA Astrophysics Data System (ADS)

Quinn, Daniel B.; Watts, Anthony; Nagle, Tony; Lentink, David

2017-03-01

Our understanding of animal flight benefits greatly from specialized wind tunnels designed for flying animals. Existing facilities can simulate laminar flow during straight, ascending and descending flight, as well as at different altitudes. However, the atmosphere in which animals fly is even more complex. Flow can be laminar and quiet at high altitudes but highly turbulent near the ground, and gusts can rapidly change wind speed. To study flight in both laminar and turbulent environments, a multi-purpose wind tunnel for studying animal and small vehicle flight was built at Stanford University. The tunnel is closed-circuit and can produce airspeeds up to 50 m s-1 in a rectangular test section that is 1.0 m wide, 0.82 m tall and 1.73 m long. Seamless honeycomb and screens in the airline together with a carefully designed contraction reduce centreline turbulence intensities to less than or equal to 0.030% at all operating speeds. A large diameter fan and specialized acoustic treatment allow the tunnel to operate at low noise levels of 76.4 dB at 20 m s-1. To simulate high turbulence, an active turbulence grid can increase turbulence intensities up to 45%. Finally, an open jet configuration enables stereo high-speed fluoroscopy for studying musculoskeletal control in turbulent flow.

Longitudinal Stability and Control Characteristics from a Flight Investigation of a Cruciform Canard Missile Configuration Having an Exposed Wing-canard Area Ratio of 16:1

NASA Technical Reports Server (NTRS)

Moul, Martin T; Wineman, Andrew R

1952-01-01

A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.

Support activities to maintain SUMS flight readiness

NASA Technical Reports Server (NTRS)

Wright, Willie

1992-01-01

The Shuttle Upper Atmosphere Mass Spectrometer (SUMS), a component experiment of the NASA Orbital Experiments Program (OEX), was flown aboard the shuttle Columbia (OV102) mounted at the forward end of the nose landing gear well with an atmospheric gas inlet system fitted to the lower fuselage (chin panel) surface. The SUMS was designed to provide atmospheric data in flow regimes inaccessible prior to the development of the Space Transportation System (STS). The experiment mission operation began about one hour prior to shuttle de-orbit entry maneuver and continued until reaching 1.6 torr (about 86 km altitude). The SUMS mass spectrometer consists of the spare unit from the Viking mission to Mars. Bendix Aerospace under contract to NASA LaRC incorporated the Viking mass spectrometer, a microprocessor based logic card, a pressurized instrument case, and the University of Texas at Dallas provided a gas inlet system into a configuration suited to interface with the shuttle Columbia. The SUMS experiment underwent static and dynamic calibration as well as vacuum maintenance before and after STS 40 shuttle flight. The SUMS flew a total of 3 times on the space shuttle Columbia. Between flights the SUMS was maintained in flight ready status. The flight data has been analyzed by the NASA LaRC Aerothermodynamics Branch. Flight data spectrum plots and reports are presented in the Appendices to the Final Technical Report for NAS1-17399.

CFIT Prevention Using Synthetic Vision

NASA Technical Reports Server (NTRS)

Arthur, Jarvis J., III; Prinzel, Lawrence J., III; Kramer, Lynda J.; Bailey, Randall E.; Parrish, Russell V.

2003-01-01

In commercial aviation, over 30-percent of all fatal accidents worldwide are categorized as Controlled Flight Into Terrain (CFIT) accidents where a fully functioning airplane is inadvertently flown into the ground, water, or an obstacle. An experiment was conducted at NASA Langley Research Center investigating the presentation of a synthetic terrain database scene to the pilot on a Primary Flight Display (PFD). The major hypothesis for the experiment is that a synthetic vision system (SVS) will improve the pilot s ability to detect and avoid a potential CFIT compared to conventional flight instrumentation. All display conditions, including the baseline, contained a Terrain Awareness and Warning System (TAWS) and Vertical Situation Display (VSD) enhanced Navigation Display (ND). Sixteen pilots each flew 22 approach - departure maneuvers in Instrument Meteorological Conditions (IMC) to the terrain challenged Eagle County Regional Airport (EGE) in Colorado. For the final run, the flight guidance cues were altered such that the departure path went into the terrain. All pilots with a SVS enhanced PFD (12 of 16 pilots) noticed and avoided the potential CFIT situation. All of the pilots who flew the anomaly with the baseline display configuration (which included a TAWS and VSD enhanced ND) had a CFIT event.

The aerodynamic challenges of the design and development of the space shuttle orbiter

NASA Technical Reports Server (NTRS)

Young, J. C.; Underwood, J. M.; Hillje, E. R.; Whitnah, A. M.; Romere, P. O.; Gamble, J. D.; Roberts, B. B.; Ware, G. M.; Scallion, W. I.; Spencer, B., Jr.

1985-01-01

The major aerodynamic design challenge at the beginning of the United States Space Transportation System (STS) research and development phase was to design a vehicle that would fly as a spacecraft during early entry and as an aircraft during the final phase of entry. The design was further complicated because the envisioned vehicle was statically unstable during a portion of the aircraft mode of operation. The second challenge was the development of preflight aerodynamic predictions with an accuracy consistent with conducting a manned flight on the initial orbital flight. A brief history of the early contractual studies is presented highlighting the technical results and management decisions influencing the aerodynamic challenges. The configuration evolution and the development of preflight aerodynamic predictions will be reviewed. The results from the first four test flights shows excellent agreement with the preflight aerodynamic predictions over the majority of the flight regimes. The only regimes showing significant disagreement is confined primarily to early entry, where prediction of the basic vehicle trim and the influence of the reaction control system jets on the flow field were found to be deficient. Postflight results are analyzed to explain these prediction deficiencies.

Analysis of structural dynamic data from Skylab. Volume 2: Skylab analytical and test model data

NASA Technical Reports Server (NTRS)

Demchak, L.; Harcrow, H.

1976-01-01

The orbital configuration test modal data, analytical test correlation modal data, and analytical flight configuration modal data are presented. Tables showing the generalized mass contributions (GMCs) for each of the thirty tests modes are given along with the two dimensional mode shape plots and tables of GMCs for the test correlated analytical modes. The two dimensional mode shape plots for the analytical modes and uncoupled and coupled modes of the orbital flight configuration at three development phases of the model are included.

14 CFR 25.123 - En route flight paths.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false En route flight paths. 25.123 Section 25... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Flight Performance § 25.123 En route flight paths. (a) For the en route configuration, the flight paths prescribed in paragraph (b) and (c) of this section must...

14 CFR 25.123 - En route flight paths.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false En route flight paths. 25.123 Section 25... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Flight Performance § 25.123 En route flight paths. (a) For the en route configuration, the flight paths prescribed in paragraph (b) and (c) of this section must...

Future Flight Decks

NASA Technical Reports Server (NTRS)

Arbuckle, P. Douglas; Abbott, Kathy H.; Abbott, Terence S.; Schutte, Paul C.

1998-01-01

The evolution of commercial transport flight deck configurations over the past 20-30 years and expected future developments are described. Key factors in the aviation environment are identified that the authors expect will significantly affect flight deck designers. One of these is the requirement for commercial aviation accident rate reduction, which is probably required if global commercial aviation is to grow as projected. Other factors include the growing incrementalism in flight deck implementation, definition of future airspace operations, and expectations of a future pilot corps that will have grown up with computers. Future flight deck developments are extrapolated from observable factors in the aviation environment, recent research results in the area of pilot-centered flight deck systems, and by considering expected advances in technology that are being driven by other than aviation requirements. The authors hypothesize that revolutionary flight deck configuration changes will be possible with development of human-centered flight deck design methodologies that take full advantage of commercial and/or entertainment-driven technologies.

Operational Overview for UAS Integration in the NAS Project Flight Test Series 3

NASA Technical Reports Server (NTRS)

Valkov, Steffi B.; Sternberg, Daniel; Marston, Michael

2018-01-01

The National Aeronautics and Space Administration Unmanned Aircraft Systems Integration in the National Airspace System Project has conducted a series of flight tests intended to support the reduction of barriers that prevent unmanned aircraft from flying without the required waivers from the Federal Aviation Administration. The 2015 Flight Test Series 3, supported two separate test configurations. The first configuration investigated the timing of Detect and Avoid alerting thresholds using a radar equipped unmanned vehicle and multiple live intruders flown at varying encounter geometries.

Linear Aerospike SR-71 Experiment (LASRE) first flight takeoff

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 takes off Oct. 31, making its first flight as part of the NASA/Rocketdyne/Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

Linear Aerospike SR-71 Experiment (LASRE) first flight view from above

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 made its successful first flight Oct. 31 as part of the NASA/Rocketdyne/Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

Linear Aerospike SR-71 Experiment (LASRE) first flight

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 successfully completed its first flight 31 October 1997 as part of the NASA/Rocketdyne/Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

Linear Aerospike SR-71 Experiment (LASRE) first flight

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 made its successful first flight Oct. 31 as part of the NASA/Rocketdyne/ Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

Integration Process for Payloads in the Fluids and Combustion Facility

NASA Technical Reports Server (NTRS)

Free, James M.; Nall, Marsha M.

2001-01-01

The Fluids and Combustion Facility (FCF) is an ISS research facility located in the United States Laboratory (US Lab), Destiny. The FCF is a multi-discipline facility that performs microgravity research primarily in fluids physics science and combustion science. This facility remains on-orbit and provides accommodations to multi-user and Principal investigator (PI) unique hardware. The FCF is designed to accommodate 15 PI's per year. In order to allow for this number of payloads per year, the FCF has developed an end-to-end analytical and physical integration process. The process includes provision of integration tools, products and interface management throughout the life of the payload. The payload is provided with a single point of contact from the facility and works with that interface from PI selection through post flight processing. The process utilizes electronic tools for creation of interface documents/agreements, storage of payload data and rollup for facility submittals to ISS. Additionally, the process provides integration to and testing with flight-like simulators prior to payload delivery to KSC. These simulators allow the payload to test in the flight configuration and perform final facility interface and science verifications. The process also provides for support to the payload from the FCF through the Payload Safety Review Panel (PSRP). Finally, the process includes support in the development of operational products and the operation of the payload on-orbit.

Mission Control Center (MCC) System Specification for the Shuttle Orbital Flight Test (OFT) Timeframe

NASA Technical Reports Server (NTRS)

1976-01-01

System specifications to be used by the mission control center (MCC) for the shuttle orbital flight test (OFT) time frame were described. The three support systems discussed are the communication interface system (CIS), the data computation complex (DCC), and the display and control system (DCS), all of which may interfere with, and share processing facilities with other applications processing supporting current MCC programs. The MCC shall provide centralized control of the space shuttle OFT from launch through orbital flight, entry, and landing until the Orbiter comes to a stop on the runway. This control shall include the functions of vehicle management in the area of hardware configuration (verification), flight planning, communication and instrumentation configuration management, trajectory, software and consumables, payloads management, flight safety, and verification of test conditions/environment.

High-speed civil transport - Advanced flight deck challenges

NASA Technical Reports Server (NTRS)

Swink, Jay R.; Goins, Richard T.

1992-01-01

This paper presents the results of a nine month study of the HSCT flight deck challenges and assessment of its benefits. Operational requirements are discussed and the most significant findings for specified advanced concepts are highlighted. These concepts are a no nose-droop configuration, a far forward cockpit location and advanced crew monitoring and control of complex systems. Results indicate that the no nose-droop configuration is critically dependent on the design and development of a safe, reliable and certifiable synthetic vision system (SVS). This configuration would cause significant weight, performance and cost penalties. A far forward cockpit configuration with a tandem seating arrangement allows either an increase in additional payload or potential downsizing of the vehicle leading to increased performance efficiency and reductions in emissions. The technologies enabling such capabilities, which provide for Category III all-weather opreations on every flight represent a benefit multiplier in a 20005 ATM network in terms of enhanced economic viability and environmental acceptability.

Evaluation of bearing configurations using the single bearing tester in liquid nitrogen

NASA Technical Reports Server (NTRS)

Jett, T.; Hall, P.; Thom, R.

1991-01-01

Various bearing configurations were tested using the Marshall Space Flight Center single bearing tester with LN2 as the cryogenic coolant. The baseline was one Rocketdyne phase one high pressure oxidizer turbopump (HPOTP) pump end 45-mm bore bearing. The bearing configurations that were tested included a Salox/M cage configuration, a silicon nitride ball configuration, an elongated cage configuration, and a Bray 601 grease configuration.

X-31 quasi-tailless flight demonstration

NASA Technical Reports Server (NTRS)

Huber, Peter; Schellenger, Harvey G.

1994-01-01

The primary objective of the quasi-tailless flight demonstration is to demonstrate the feasibility of using thrust vectoring for directional control of an unstable aircraft. By using this low-cost, low-risk approach it is possible to get information about required thrust vector control power and deflection rates from an inflight experiment as well as insight in low-power thrust vectoring issues. The quasi-tailless flight demonstration series with the X-31 began in March 1994. The demonstration flight condition was Mach 1.2 at 37,500 feet. A series of basic flying quality maneuvers, doublets, bank to bank rolls, and wind-up-turns have been performed with a simulated 100% vertical tail reduction. Flight test and supporting simulation demonstrated that the quasi-tailless approach is effective in representing the reduced stability of tailless configurations. The flights also demonstrated that thrust vectoring could be effectively used to stabilize a directionally unstable configuration and provide control power for maneuver coordination.

Experimental Aerodynamic Characteristics of the Pegasus Air-Launched Booster and Comparisons with Predicted and Flight Results

NASA Technical Reports Server (NTRS)

Rhode, M. N.; Engelund, Walter C.; Mendenhall, Michael R.

1995-01-01

Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for the Pegasus and Pegasus XL configurations over a Mach number range from 1.6 to 6 and angles of attack from -4 to +24 degrees. Angle of sideslip was varied from -6 to +6 degrees, and control surfaces were deflected to obtain elevon, aileron, and rudder effectiveness. Experimental data for the Pegasus configuration are compared with engineering code predictions performed by Nielsen Engineering & Research, Inc. (NEAR) in the aerodynamic design of the Pegasus vehicle, and with results from the Aerodynamic Preliminary Analysis System (APAS) code. Comparisons of experimental results are also made with longitudinal flight data from Flight #2 of the Pegasus vehicle. Results show that the longitudinal aerodynamic characteristics of the Pegasus and Pegasus XL configurations are similar, having the same lift-curve slope and drag levels across the Mach number range. Both configurations are longitudinally stable, with stability decreasing towards neutral levels as Mach number increases. Directional stability is negative at moderate to high angles of attack due to separated flow over the vertical tail. Dihedral effect is positive for both configurations, but is reduced 30-50 percent for the Pegasus XL configuration because of the horizontal tail anhedral. Predicted longitudinal characteristics and both longitudinal and lateral-directional control effectiveness are generally in good agreement with experiment. Due to the complex leeside flowfield, lateral-directional characteristics are not as well predicted by the engineering codes. Experiment and flight data are in good agreement across the Mach number range.

Manned remote work station development article. Volume 1, book 1: Flight article requirements. Appendix A: Mission requirements

NASA Technical Reports Server (NTRS)

1979-01-01

The requirements for several configurations of flight articles are presented. These requirements provide the basis to design manned remote work station development test articles and establish tests and simulation objectives for the resolution of development issues. Mission system and subsystem requirements for four MRWS configurations included: open cherry picker; closed cherry picker; crane turret; and free flyer.

Space Launch System Co-Manifested Payload Options for Habitation

NASA Technical Reports Server (NTRS)

Smitherman, David

2015-01-01

The Space Launch System (SLS) has a co-manifested payload capability that will grow over time as the launch vehicle matures and planned upgrades are implemented. The final configuration is planned to be capable of inserting a payload greater than 10 metric tons (mt) into a trans-lunar injection trajectory along with the crew in the Orion capsule and its service module. The co-manifested payload is located below the Orion and its service module in a 10 m high fairing similar to the way the Saturn launch vehicle carried the lunar lander below the Apollo command and service modules. Various approaches that utilize this comanifested payload capability to build up infrastructure in deep space have been explored in support of future asteroid, lunar, and Mars mission scenarios. This paper reports on the findings of the Advanced Concepts Office study team at NASA Marshall Space Flight Center (MSFC) working with the Advanced Exploration Systems Program on the Exploration Augmentation Module Project. It includes some of the possible options for habitation in the co-manifested payload volume of the SLS. Findings include a set of module designs that can be developed in 10 mt increments to support these co-manifested payload missions along with a comparison of this approach to a large-module payload flight configuration for the SLS.

Subsystem Analysis/Optimization for the X-34 Main Propulsion System

NASA Technical Reports Server (NTRS)

McDonald, J. P.; Hedayat, A.; Brown, T. M.; Knight, K. C.; Champion, R. H., Jr.

1998-01-01

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System (MPS) supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major MPS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. Analyses related to optimal MPS subsystem design are reviewed in this paper. A pressurization system trade weighs maintenance/reliability concerns against those for safety in a comparison of designs using pressure regulators versus orifices to control pressurant flow. A propellant dump/feed system analysis weighs the issues of maximum allowable vehicle landing weight, trajectory, and MPS complexity to arrive at a final configuration for propellant dump/feed systems.

Orbiter Docking System/Spacelab-Mir Module in Atlantis

NASA Technical Reports Server (NTRS)

1995-01-01

The STS-71 mission payload is in its final flight configuration after integration into the payload bay of the Space Shuttle orbiter Atlantis and prior to payload bay door closing and rollover of the spaceplane from Orbiter Processing Facility Bay 3 to the Vehicle Assembly Building. In the foreground is the Orbiter Docking System (ODS) that is topped with the red Russian- built Androgynous Peripheral Docking System (APDS). During the 11-day mission, the APDS will lock together with a similar system on the Russian Mir Space Station so that the two spacecraft can remain docked together for four days. The ODS features an airlock that will provide access to and from both the Mir and orbiter for the U.S. and Russian flight crews. A Spacelab transfer tunnel runs from the ODS to the Spacelab-Mir module, where joint U.S. medical experiments will be conducted during the 11-day spaceflight.

Saturn Apollo Program

NASA Image and Video Library

1968-12-19

Pictured from left to right, the Apollo 9 astronauts, James A. McDivitt, David R. Scott, and Russell L. Schweickart, pause in front of the Apollo/Saturn V space vehicle that would launch the Apollo 8 crew. The launch of the Apollo 9 (Saturn V launch vehicle, SA-504) took place on March 3, 1968. The Apollo 9 spacecraft, in the lunar mission configuration, was tested in Earth orbit. The mission was designed to rehearse all the steps and reproduce all the events of the Apollo 11 mission with the exception of the lunar touchdown, stay, and liftoff. The command and service modules, and the lunar module were used in flight procedures identical to those that would later take similar vehicles to the Moon, and a landing. The flight mechanics, mission support systems, communications, and recording of data were tested in a final round of verification. Astronauts Scott and Schweickart conducted Extravehicular Activity during this mission.

International Space Station (ISS)

NASA Image and Video Library

2005-06-09

The STS-121 patch depicts the Space Shuttle docked with the International Space Station (ISS) in the foreground, overlaying the astronaut symbol with three gold columns and a gold star. The ISS is shown in the configuration that it was during the STS-121 mission. The background shows the nighttime Earth with a dawn breaking over the horizon. STS-121, ISS mission ULF1.1, was the final Shuttle Return to Flight test mission. This utilization and logistics flight delivered a multipurpose logistics module (MPLM) to the ISS with several thousand pounds of new supplies and experiments. In addition, some new orbital replacement units (ORUs) were delivered and stowed externally on the ISS on a special pallet. These ORUs are spares for critical machinery located on the outside of the ISS. During this mission the crew also carried out testing of Shuttle inspection and repair hardware, as well as evaluated operational techniques and concepts for conducting on-orbit inspection and repair.

ACAS-Xu Initial Self-Separation Flight Tests

NASA Technical Reports Server (NTRS)

Marston, Mike; Baca, Gabe

2015-01-01

The purpose of this flight test report is to document and report the details of the ACAS Xu (Airborne Collision Avoidance System For Unmanned Aircraft) / Self-Separation flight test series performed at Edwards AFB from November to December of 2014. Included in this document are details about participating aircraft, aircrew, mission crew, system configurations, flight data, flight execution, flight summary, test results, and lessons learned.

A comparison of theory and experiment for coupled rotor body stability of a bearingless rotor model in hover and forward flight

NASA Technical Reports Server (NTRS)

Mirick, Paul H.

1988-01-01

Seven cases were selected for correlation from a 1/5.86 Froude-scale experiment that examined several rotor designs which were being considered for full-scale flight testing as part of the Bearingless Main Rotor (BMR) program. The model rotor hub used in these tests consisted of back-to-back C-beams as flexbeam elements with a torque tube for pitch control. The first four cases selected from the experiment were hover tests which examined the effects on rotor stability of variations in hub-to-flexbeam coning, hub-to-flexbeam pitch, flexbeam-to-blade coning, and flexbeam-to-blade pitch. The final three cases were selected from the forward flight tests of optimum rotor configuration as defined during the hover test. The selected cases examined the effects of variations in forward speed, rotor speed, and shaft angle. Analytical results from Bell Helicopter Textron, Boeing Vertol, Sikorsky Aircraft, and the U.S. Army Aeromechanics Laboratory were compared with the data and the correlations ranged from poor-to-fair to fair-to-good.

A dual-processor multi-frequency implementation of the FINDS algorithm

NASA Technical Reports Server (NTRS)

Godiwala, Pankaj M.; Caglayan, Alper K.

1987-01-01

This report presents a parallel processing implementation of the FINDS (Fault Inferring Nonlinear Detection System) algorithm on a dual processor configured target flight computer. First, a filter initialization scheme is presented which allows the no-fail filter (NFF) states to be initialized using the first iteration of the flight data. A modified failure isolation strategy, compatible with the new failure detection strategy reported earlier, is discussed and the performance of the new FDI algorithm is analyzed using flight recorded data from the NASA ATOPS B-737 aircraft in a Microwave Landing System (MLS) environment. The results show that low level MLS, IMU, and IAS sensor failures are detected and isolated instantaneously, while accelerometer and rate gyro failures continue to take comparatively longer to detect and isolate. The parallel implementation is accomplished by partitioning the FINDS algorithm into two parts: one based on the translational dynamics and the other based on the rotational kinematics. Finally, a multi-rate implementation of the algorithm is presented yielding significantly low execution times with acceptable estimation and FDI performance.

Aeroacoustics of advanced propellers

NASA Technical Reports Server (NTRS)

Groeneweg, John F.

1990-01-01

The aeroacoustics of advanced, high speed propellers (propfans) are reviewed from the perspective of NASA research conducted in support of the Advanced Turboprop Program. Aerodynamic and acoustic components of prediction methods for near and far field noise are summarized for both single and counterrotation propellers in uninstalled and configurations. Experimental results from tests at both takeoff/approach and cruise conditions are reviewed with emphasis on: (1) single and counterrotation model tests in the NASA Lewis 9 by 15 (low speed) and 8 by 6 (high speed) wind tunnels, and (2) full scale flight tests of a 9 ft (2.74 m) diameter single rotation wing mounted tractor and a 11.7 ft (3.57 m) diameter counterrotation aft mounted pusher propeller. Comparisons of model data projected to flight with full scale flight data show good agreement validating the scale model wind tunnel approach. Likewise, comparisons of measured and predicted noise level show excellent agreement for both single and counterrotation propellers. Progress in describing angle of attack and installation effects is also summarized. Finally, the aeroacoustic issues associated with ducted propellers (very high bypass fans) are discussed.

Air launch wireless sensor nodes (ALSN) for battle damage assessment (BDA)

NASA Astrophysics Data System (ADS)

Back, Jason M.; Beck, Steven D.; Frank, Mark A.; Hoenes, Eric

2006-05-01

This paper summarizes the Defense Threat Reduction Agency (DTRA) sponsored development and demonstration of an Air Launched Sensor Node (ALSN) system designed to fill DTRA's immediate need to support the Global Strike requirement of weapon-borne deliverable sensors for Battle Damage Assessment (BDA). Unattended ground sensors were integrated into a CBU-103 Tactical Munitions Dispenser (TMD), and flight test demonstrated with the 46 th Test Wing at Eglin AFB, FL. The objectives of the ALSN program were to repackage an existing multi-sensor node system to conform to the payload envelope and deployment configuration design; to integrate this payload into the CBU-103 TMD; and to conduct a combined payload flight test demonstration. The final sensor node included multiple sensors a microphone, a geophone, and multiple directional Passive Infrared (PIR) detectors with processing electronics, a low power wireless communications 802.15.4 mesh network, GPS (Global Positioning System), and power integrated into a form-fit BLU-97 munitions deployable package. This paper will present and discuss the flight test, results, and ALSN performance.

Investigation of aeroelastic stability phenomena of a helicopter by in-flight shake test

NASA Technical Reports Server (NTRS)

Miao, W. L.; Edwards, T.; Brandt, D. E.

1976-01-01

The analytical capability of the helicopter stability program is discussed. The parameters which are found to be critical to the air resonance characteristics of the soft in-plane hingeless rotor systems are detailed. A summary of two model test programs, a 1/13.8 Froude-scaled BO-105 model and a 1.67 meter (5.5 foot) diameter Froude-scaled YUH-61A model, are presented with emphasis on the selection of the final parameters which were incorporated in the full scale YUH-61A helicopter. Model test data for this configuration are shown. The actual test results of the YUH-61A air resonance in-flight shake test stability are presented. Included are a concise description of the test setup, which employs the Grumman Automated Telemetry System (ATS), the test technique for recording in-flight stability, and the test procedure used to demonstrate favorable stability characteristics with no in-plane damping augmentation (lag damper removed). The data illustrating the stability trend of air resonance with forward speed and the stability trend of ground resonance for percent airborne are presented.

Design, Analysis and Qualification of Elevon for Reusable Launch Vehicle

NASA Astrophysics Data System (ADS)

Tiwari, S. B.; Suresh, R.; Krishnadasan, C. K.

2017-12-01

Reusable launch vehicle technology demonstrator is configured as a winged body vehicle, designed to fly in hypersonic, supersonic and subsonic regimes. The vehicle will be boosted to hypersonic speeds after which the winged body separates and descends using aerodynamic control. The aerodynamic control is achieved using the control surfaces mainly the rudder and the elevon. Elevons are deflected for pitch and roll control of the vehicle at various flight conditions. Elevons are subjected to aerodynamic, thermal and inertial loads during the flight. This paper gives details about the configuration, design, qualification and flight validation of elevon for Reusable Launch Vehicle.

Recent Loads Calibration Experience With a Delta Wing Airplane

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.; Kuhl, Albert E.

1977-01-01

Aircraft which are designed for supersonic and hypersonic flight are evolving with delta wing configurations. An integral part of the evolution of all new aircraft is the flight test phase. Included in the flight test phase is an effort to identify and evaluate the loads environment of the aircraft. The most effective way of examining the loads environment is to utilize calibrated strain gages to provide load magnitudes. Using strain gage data to accomplish this has turned out to be anything but a straightforward task. The delta wing configuration has turned out to be a very difficult type of wing structure to calibrate. Elevated structural temperatures result in thermal effects which contaminate strain gage data being used to deduce flight loads. The concept of thermally calibrating a strain gage system is an approach to solving this problem. This paper will address how these problems were approached on a program directed toward measuring loads on the wing of a large, flexible supersonic aircraft. Structural configurations typical of high-speed delta wing aircraft will be examined. The temperature environment will be examined to see how it induces thermal stresses which subsequently cause errors in loads equations used to deduce the flight loads.

Test results of flight guidance for fuel conservative descents in a time-based metered air traffic environment. [terminal configured vehicle

NASA Technical Reports Server (NTRS)

Knox, C. E.; Person, L. H., Jr.

1981-01-01

The NASA developed, implemented, and flight tested a flight management algorithm designed to improve the accuracy of delivering an airplane in a fuel-conservative manner to a metering fix at a time designated by air traffic control. This algorithm provides a 3D path with time control (4D) for the TCV B-737 airplane to make an idle-thrust, clean configured (landing gear up, flaps zero, and speed brakes retracted) descent to arrive at the metering fix at a predetermined time, altitude, and airspeed. The descent path is calculated for a constant Mach/airspeed schedule from linear approximations of airplane performance with considerations given for gross weight, wind, and nonstandard pressure and temperature effects. The flight management descent algorithms are described and flight test results are presented.

Evaluation of helicopter noise due to b blade-vortex interaction for five tip configurations. [conducted in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Hoad, D. R.

1979-01-01

The effect of tip shape modification on blade vortex interaction induced helicopter blade slap noise was investigated. Simulated flight and descent velocities which have been shown to produce blade slap were tested. Aerodynamic performance parameters of the rotor system were monitored to ensure properly matched flight conditions among the tip shapes. The tunnel was operated in the open throat configuration with treatment to improve the acoustic characteristics of the test chamber. Four promising tips were used along with a standard square tip as a baseline configuration. A detailed acoustic evaluation on the same rotor system of the relative applicability of the various tip configurations for blade slap noise reduction is provided.

RICA: a reliable and image configurable arena for cyborg bumblebee based on CAN bus.

PubMed

Gong, Fan; Zheng, Nenggan; Xue, Lei; Xu, Kedi; Zheng, Xiaoxiang

2014-01-01

In this paper, we designed a reliable and image configurable flight arena, RICA, for developing cyborg bumblebees. To meet the spatial and temporal requirements of bumblebees, the Controller Area Network (CAN) bus is adopted to interconnect the LED display modules to ensure the reliability and real-time performance of the arena system. Easily-configurable interfaces on a desktop computer implemented by python scripts are provided to transmit the visual patterns to the LED distributor online and configure RICA dynamically. The new arena system will be a power tool to investigate the quantitative relationship between the visual inputs and induced flight behaviors and also will be helpful to the visual-motor research in other related fields.

System Design and Nonlinear State-Dependent Riccati Equation Control of an Autonomous Y-4 Tilt-Rotor Aerobot for Martian Exploration

NASA Astrophysics Data System (ADS)

Collins, Nathan Scott

Surrey Space Centre (SSC) has been working on an autonomous fixed-wing all-electric vertical take-off and landing (VTOL) aerobot for the exploration of Mars for several years. SSC's previous designs have incorporated separate vertical lift and horizontal pusher rotors as well as a mono tilt-rotor configuration. The Martian aerobot's novel Y-4 tilt-rotor (Y4TR) design is a combination of two previous SSC designs and a step forward for planetary aerobots. The aerobot will fly as a Y4 multi-rotor during vertical flight and as a conventional flying wing during horizontal flight. The more robust Y4TR configuration utilizes two large fixed coaxial counter rotating rotors and two small tilt-rotors for vertical takeoff. The front tilt-rotors rotate during transition flight into the main horizontal flight configuration. The aerobot is a blended wing design with the wings using the "Zagi 10" airfoil blended to a center cover for the coaxial rotors. The open source design and analysis programs XROTOR, CROTOR, Q-BLADE, XFLR5, and OpenVSP were used to design and model the aerobot's four rotors and body. The baseline mission of the Y4TR remains the same as previously reported and will investigate the Isidis Planitia region on Mars over a month long period using optical sensors during flight and a surface science package when landed. During flight operations the aerobot will take off vertically, transition to horizontal flight, fly for around an hour, transition back to vertical flight, and land vertically. The flight missions will take place close to local noon to maximize power production via solar cells during flight. A nonlinear six degree of freedom (6DoF) dynamic model incorporating aerodynamic models of the aerobot's body and rotors has been developed to model the vertical, transition, and horizontal phases of flight. A nonlinear State-Dependent Riccati Equation (SDRE) controller has been developed for each of these flight phases. The nonlinear dynamic model was transformed into a pseudo-linear form based on the states and implemented in the SDRE controller. During transition flight the aerobot is over actuated and the weighted least squares (WLS) method is used for allocation of control effectors. Simulations of the aerobot flying in different configurations were performed to verify the performance of the SDRE controllers, including hover, transition, horizontal flight, altitude changes, and landing scenarios. Results from the simulations show the SDRE controller is a viable option for controlling the novel Y4TR Martian Aerobot.

Isomer-delayed gamma-ray spectroscopy of neutron-rich 166Tb

DOE PAGES

Gurgi, L. A.; Regan, P. H.; Söderström, P. -A.; ...

2017-09-13

Here, this short paper presents the identification of a metastable, isomeric-state decay in the neutron-rich odd-odd, prolate-deformed nucleus 166Tb. The nucleus of interest was formed using the in-flight fission of a 345 MeV per nucleon 238U primary beam at the RIBF facility, RIKEN, Japan. Gamma-ray transitions decaying from the observed isomeric states in 166Tb were identified using the EURICA gamma-ray spectrometer, positioned at the final focus of the BigRIPS fragments separator. The current work identifies a single discrete gamma-ray transition of energy 119 keV which de-excites an isomeric state in 166Tb with a measured half-life of 3.5(4) μs. The multipolaritymore » assignment for this transition is an electric dipole and is made on the basis internal conversion and decay lifetime arguments. Possible two quasi-particle Nilsson configurations for the initial and final states which are linked by this transition in 166Tb are made on the basis of comparison with Blocked BCS Nilsson calculations, with the predicted ground state configuration for this nucleus arising from the coupling of the v(1-/2)[521] and π(3+/2) Nilsson orbitals.« less

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gurgi, L. A.; Regan, P. H.; Söderström, P. -A.

Here, this short paper presents the identification of a metastable, isomeric-state decay in the neutron-rich odd-odd, prolate-deformed nucleus 166Tb. The nucleus of interest was formed using the in-flight fission of a 345 MeV per nucleon 238U primary beam at the RIBF facility, RIKEN, Japan. Gamma-ray transitions decaying from the observed isomeric states in 166Tb were identified using the EURICA gamma-ray spectrometer, positioned at the final focus of the BigRIPS fragments separator. The current work identifies a single discrete gamma-ray transition of energy 119 keV which de-excites an isomeric state in 166Tb with a measured half-life of 3.5(4) μs. The multipolaritymore » assignment for this transition is an electric dipole and is made on the basis internal conversion and decay lifetime arguments. Possible two quasi-particle Nilsson configurations for the initial and final states which are linked by this transition in 166Tb are made on the basis of comparison with Blocked BCS Nilsson calculations, with the predicted ground state configuration for this nucleus arising from the coupling of the v(1-/2)[521] and π(3+/2) Nilsson orbitals.« less

Isomer-delayed gamma-ray spectroscopy of neutron-rich 166Tb

NASA Astrophysics Data System (ADS)

Gurgi, L. A.; Regan, P. H.; Söderström, P.-A.; Watanabe, H.; Walker, P. M.; Podolyák, Zs.; Nishimura, S.; Berry, T. A.; Doornenbal, P.; Lorusso, G.; Isobe, T.; Baba, H.; Xu, Z. Y.; Sakurai, H.; Sumikama, T.; Catford, W. N.; Bruce, A. M.; Browne, F.; Lane, G. J.; Kondev, F. G.; Odahara, A.; Wu, J.; Liu, H. L.; Xu, F. R.; Korkulu, Z.; Lee, P.; Liu, J. J.; Phong, V. H.; Yagi, A.; Zhang, G. X.; Alharbi, T.; Carroll, R. J.; Chae, K. Y.; Dombradi, Zs.; Estrade, A.; Fukuda, N.; Griffin, C.; Ideguchi, E.; Inabe, N.; Kanaoka, H.; Kojouharov, I.; Kubo, T.; Kubono, S.; Kurz, N.; Kuti, I.; Lalkovski, S.; Lee, E. J.; Lee, C. S.; Lotay, G.; Moon, C. B.; Nishizuka, I.; Nita, C. R.; Patel, Z.; Roberts, O. J.; Schaffner, H.; Shand, C. M.; Suzuki, H.; Takeda, H.; Terashima, S.; Vajta, Zs.; Kanaya, S.; Valiente-Dobòn, J. J.

2017-09-01

This short paper presents the identification of a metastable, isomeric-state decay in the neutron-rich odd-odd, prolate-deformed nucleus 166Tb. The nucleus of interest was formed using the in-flight fission of a 345 MeV per nucleon 238U primary beam at the RIBF facility, RIKEN, Japan. Gamma-ray transitions decaying from the observed isomeric states in 166Tb were identified using the EURICA gamma-ray spectrometer, positioned at the final focus of the BigRIPS fragments separator. The current work identifies a single discrete gamma-ray transition of energy 119 keV which de-excites an isomeric state in 166Tb with a measured half-life of 3.5(4) μs. The multipolarity assignment for this transition is an electric dipole and is made on the basis internal conversion and decay lifetime arguments. Possible two quasi-particle Nilsson configurations for the initial and final states which are linked by this transition in 166Tb are made on the basis of comparison with Blocked BCS Nilsson calculations, with the predicted ground state configuration for this nucleus arising from the coupling of the v(1-/2)?[521] and ? π(3+/2) Nilsson orbitals.

An Overview of an Experimental Demonstration Aerotow Program

NASA Technical Reports Server (NTRS)

Murray, James E.; Bowers, Albion H.; Lokos, William A.; Peters, Todd L.; Gera, Joseph

1998-01-01

An overview of an experimental demonstration of aerotowing a delta-wing airplane with low-aspect ratio and relatively high wing loading is presented. Aerotowing of future space launch configurations is a new concept, and the objective of the work described herein is to demonstrate the aerotow operation using an airplane configuration similar to conceptual space launch vehicles. Background information on the use of aerotow for a space launch vehicle is presented, and the aerotow system used in this demonstration is described. The ground tests, analytical studies, and flight planning used to predict system behavior and to enhance flight safety are detailed. The instrumentation suite and flight test maneuvers flown are discussed, preliminary performance is assessed, and flight test results are compared with the preflight predictions.

KENNEDY SPACE CENTER, FLA. - STS-82 crew members and workers at KSC's Vertical Processing Facility get a final look at the Near Infrared Camera and Multi-Object Spectrometer (NICMOS) in its flight configuration for the STS-82 mission. The crew is participating in the Crew Equipment Integration Test (CEIT). NICMOS is one of two new scientific instruments that will replace two outdated instruments on the Hubble Space Telescope (HST). NICMOS will provide HST with the capability for infrared imaging and spectroscopic observations of astronomical targets. The refrigerator-sized NICMOS also is HST's first cryogenic instrument - its sensitive infrared detectors must operate at very cold temperatures of minus 355 degrees Fahrenheit or 58 degrees Kelvin. NICMOS will be installed in Hubble during STS-82, the second Hubble Space Telescope servicing mission. Liftoff is scheduled Feb. 11 aboard Discovery with a crew of seven.

NASA Image and Video Library

1997-01-22

KENNEDY SPACE CENTER, FLA. - STS-82 crew members and workers at KSC's Vertical Processing Facility get a final look at the Near Infrared Camera and Multi-Object Spectrometer (NICMOS) in its flight configuration for the STS-82 mission. The crew is participating in the Crew Equipment Integration Test (CEIT). NICMOS is one of two new scientific instruments that will replace two outdated instruments on the Hubble Space Telescope (HST). NICMOS will provide HST with the capability for infrared imaging and spectroscopic observations of astronomical targets. The refrigerator-sized NICMOS also is HST's first cryogenic instrument - its sensitive infrared detectors must operate at very cold temperatures of minus 355 degrees Fahrenheit or 58 degrees Kelvin. NICMOS will be installed in Hubble during STS-82, the second Hubble Space Telescope servicing mission. Liftoff is scheduled Feb. 11 aboard Discovery with a crew of seven.

A new low-turbulence wind tunnel for animal and small vehicle flight experiments

PubMed Central

Watts, Anthony; Nagle, Tony; Lentink, David

2017-01-01

Our understanding of animal flight benefits greatly from specialized wind tunnels designed for flying animals. Existing facilities can simulate laminar flow during straight, ascending and descending flight, as well as at different altitudes. However, the atmosphere in which animals fly is even more complex. Flow can be laminar and quiet at high altitudes but highly turbulent near the ground, and gusts can rapidly change wind speed. To study flight in both laminar and turbulent environments, a multi-purpose wind tunnel for studying animal and small vehicle flight was built at Stanford University. The tunnel is closed-circuit and can produce airspeeds up to 50 m s−1 in a rectangular test section that is 1.0 m wide, 0.82 m tall and 1.73 m long. Seamless honeycomb and screens in the airline together with a carefully designed contraction reduce centreline turbulence intensities to less than or equal to 0.030% at all operating speeds. A large diameter fan and specialized acoustic treatment allow the tunnel to operate at low noise levels of 76.4 dB at 20 m s−1. To simulate high turbulence, an active turbulence grid can increase turbulence intensities up to 45%. Finally, an open jet configuration enables stereo high-speed fluoroscopy for studying musculoskeletal control in turbulent flow. PMID:28405384

Launch, Low-Speed, and Landing Characteristics Determined from the First Flight of the North American X-15 Research Airplane

NASA Technical Reports Server (NTRS)

Finch, Thomas W.; Matranga, Gene J.

1959-01-01

The first flight of the North American X-15 research airplane was made on June 8, 1959. This was accomplished after completion of a series of captive flights with the X-15 attached to the B-52 carrier airplane to demonstrate the aerodynamic and systems compatibility of the X-15//B-52 combination and the X-15 subsystem operation. This flight was planned as a glide flight so that the pilot need not be concerned with the propulsion system. Discussions of the launch, low-speed maneuvering, and landing characteristics are presented, and the results are compared with predictions from preflight studies. The launch characteristics were generally satisfactory, and the X-15 vertical tail adequately cleared the B-52 wing cutout. The actual landing pattern and landing characteristics compared favorably with predictions, and the recommended landing technique of lowering the flaps and landing gear at a low altitude appears to be a satisfactory method of landing the X-15 airplane. There was a quantitative correlation between flight-measured and predicted lift-drag-ratio characteristics in the clean configuration and a qualitative correlation in the landing configuration. A longitudinal-controllability problem, which became severe in the landing configuration, was evident throughout the flight and, apparently, was aggravated by the sensitivity of the side-located control stick. In the low-to-moderate angle-of-attack range covered, the longitudinal and directional stability were indicated to be adequate.

Post-flight BET products for the 2nd discovery entry, STS-19 (51-A)

NASA Technical Reports Server (NTRS)

Kelly, G. M.; Mcconnell, J. G.; Heck, M. L.; Troutman, P. A.; Waters, L. A.; Findlay, J. T.

1985-01-01

The post-flight products for the second Discovery flight, STS-19 (51-A), are summarized. The inertial best estimate trajectory (BET), BT19D19/UN=169750N, was developed using spacecraft dynamic measurements from Inertial Measurement Unit 2 (IMU2) in conjunction with the best tracking coverage available for any of the earlier Shuttle entries. As a consequence of the latter, an anchor epoch was selected which conforms to an initial altitude of greater than a million feet. The Extended BET, ST19BET/UN=274885C, incorporated the previously mentioned inertial reconstructed state information and the Langley Atmospheric Information Retrieval System (LAIRS) atmosphere, ST19MET/UN=712662N, with some minor exceptions. Primary and back-up AEROBET reels are NK0165 and NK0201, respectively. This product was only developed over the lowermost 360 kft altitude range due to atmosphere problems but this relates to altitudes well above meaningful signal in the IMUs. Summary results generated from the AEROBET for this flight are presented with meaningful configuration and statistical comparisons from the previous thirteen flights. Modified maximum likelihood estimation (MMLE) files were generated based on IMU2 and the Rate Gyro Assembly/Accelerometer Assembly (RGA/AA), respectively. Appendices attached define spacecraft and physical constants utilized, show plots of the final tracking data residuals from the post-flight fit, list relevant parameters from the BET at a two second spacing, and retain for archival purpose all relevant input and output tapes and files generated.

Final postflight hardware evaluation report RSRM-28 (STS-53)

NASA Technical Reports Server (NTRS)

Starrett, William David, Jr.

1993-01-01

The final report for the Clearfield disassembly evaluation and a continuation of the KSC postflight assessment for the RSRM-28 (STS-53) RSRM flight set is presented. All observed hardware conditions were documented on PFOR's and are included in Appendices A through C. Appendices D and E contain the measurements and safety factor data for the nozzle and insulation components. This report, along with the KSC Ten-Day Postflight Hardware Evaluation Report (TWR-64215), represents a summary of the RSRM-28 hardware evaluation. The as-flown hardware configuration is documented in TWR-63638. Disassembly evaluation photograph numbers are logged in TWA-1989. The RSRM-28 flight set disassembly evaluations described were performed at the RSRM Refurbishment Facility in Clearfield, Utah. The final factory joint demate occurred on July 15, 1993. Additional time was required to perform the evaluation of the stiffener rings per special issue 4.1.5.2 because of the washout schedule. The release of this report was after completion of all special issues per program management direction. Detailed evaluations were performed in accordance with the Clearfield PEEP, TWR-50051, Revision A. All observations were compared against limits that are also defined in the PEEP. These limits outline the criteria for categorizing the observations as acceptable, reportable, or critical. Hardware conditions that were unexpected and/or determined to be reportable or critical were evaluated by the applicable team and tracked through the PFAR system.

150 Passenger Commercial Aircraft

NASA Technical Reports Server (NTRS)

Bucovsky, Adrian; Romli, Fairuz I.; Rupp, Jessica

2002-01-01

It has been projected that the need for a short-range mid-sized, aircraft is increasing. The future strategy to decrease long-haul flights will increase the demand for short-haul flights. Since passengers prefer to meet their destinations quickly, airlines will increase the frequency of flights, which will reduce the passenger load on the aircraft. If a point-to-point flight is not possible, passengers will prefer only a one-stop short connecting flight to their final destination. A 150-passenger aircraft is an ideal vehicle for these situations. It is mid-sized aircraft and has a range of 3000 nautical miles. This type of aircraft would market U.S. domestic flights or inter-European flight routes. The objective of the design of the 150-passenger aircraft is to minimize fuel consumption. The configuration of the aircraft must be optimized. This aircraft must meet CO2 and NOx emissions standards with minimal acquisition price and operating costs. This report contains all the work that has been performed for the completion of the design of a 150 passenger commercial aircraft. The methodology used is the Technology Identification, Evaluation, and Selection (TIES) developed at Georgia Tech Aerospace Systems Design laboratory (ASDL). This is an eight-step conceptual design process to evaluate the probability of meeting the design constraints. This methodology also allows for the evaluation of new technologies to be implemented into the design. The TIES process begins with defining the problem with a need established and a market targeted. With the customer requirements set and the target values established, a baseline concept is created. Next, the design space is explored to determine the feasibility and viability of the baseline aircraft configuration. If the design is neither feasible nor viable, new technologies can be implemented to open up the feasible design space and allow for a plausible solution. After the new technologies are identified, they must be evaluated to determine the physical compatibility of integrating multiple technologies and then the impact on the design, both improvements and degradations, must be determined. These technologies are assessed deterministically. Again, Response Surface Equations (RSEs) are developed to allow for a full factorial evaluation of the combinations of the technologies. The best combination of technologies is selected and then the design space is again reevaluated for feasibility and viability.

UAV State Estimation Modeling Techniques in AHRS

NASA Astrophysics Data System (ADS)

Razali, Shikin; Zhahir, Amzari

2017-11-01

Autonomous unmanned aerial vehicle (UAV) system is depending on state estimation feedback to control flight operation. Estimation on the correct state improves navigation accuracy and achieves flight mission safely. One of the sensors configuration used in UAV state is Attitude Heading and Reference System (AHRS) with application of Extended Kalman Filter (EKF) or feedback controller. The results of these two different techniques in estimating UAV states in AHRS configuration are displayed through position and attitude graphs.

Modeling Pilot Behavior for Assessing Integrated Alert and Notification Systems on Flight Decks

NASA Technical Reports Server (NTRS)

Cover, Mathew; Schnell, Thomas

2010-01-01

Numerous new flight deck configurations for caution, warning, and alerts can be conceived; yet testing them with human-in-the-Ioop experiments to evaluate each one would not be practical. New sensors, instruments, and displays are being put into cockpits every day and this is particularly true as we enter the dawn of the Next Generation Air Transportation System (NextGen). By modeling pilot behavior in a computer simulation, an unlimited number of unique caution, warning, and alert configurations can be evaluated 24/7 by a computer. These computer simulations can then identify the most promising candidate formats to further evaluate in higher fidelity, but more costly, Human-in-the-Ioop (HITL) simulations. Evaluations using batch simulations with human performance models saves time, money, and enables a broader consideration of possible caution, warning, and alerting configurations for future flight decks.

Directional Stability of Towed Airplanes

NASA Technical Reports Server (NTRS)

Soehne, W.

1956-01-01

So far, very careful investigations have been made regarding the flight properties, in particular the static and dynamic stability, of engine-propelled aircraft and of untowed gliders. In contrast, almost no investigations exist regarding the stability of airplanes towed by a towline. Thus, the following report will aim at investigating the directional stability of the towed airplane and, particularly, at determining what parameters of the flight attitude and what configuration properties affect the stability. The most important parameters of the flight attitude are the dynamic pressure, the aerodynamic coefficients of the flight attitude, and the climbing angle. Among the configuration properties, the following exert the greatest influence on the stability: the tow-cable length, the tow-cable attachment point, the ratio of the wing loadings of the towing and the towed airplanes, the moments of inertia, and the wing dihedral of the towed airplane. In addition, the size and shape of the towed airplane vertical tail, the vertical tail length, and the fuselage configuration are decisive factors in determining the yawing moment and side force due to sideslip, respectively.

Terminal configured vehicle program: Test facilities guide

NASA Technical Reports Server (NTRS)

1980-01-01

The terminal configured vehicle (TCV) program was established to conduct research and to develop and evaluate aircraft and flight management system technology concepts that will benefit conventional take off and landing operations in the terminal area. Emphasis is placed on the development of operating methods for the highly automated environment anticipated in the future. The program involves analyses, simulation, and flight experiments. Flight experiments are conducted using a modified Boeing 737 airplane equipped with highly flexible display and control equipment and an aft flight deck for research purposes. The experimental systems of the Boeing 737 are described including the flight control computer systems, the navigation/guidance system, the control and command panel, and the electronic display system. The ground based facilities used in the program are described including the visual motion simulator, the fixed base simulator, the verification and validation laboratory, and the radio frequency anechoic facility.

Summary of flight tests to determine the spin and controllability characteristics of a remotely piloted, large-scale (3/8) fighter airplane model

NASA Technical Reports Server (NTRS)

Holleman, E. C.

1976-01-01

An unpowered, large, dynamically scaled airplane model was test flown by remote pilot to investigate the stability and controllability of the configuration at high angles of attack. The configuration proved to be departure/spin resistant; however, spins were obtained by using techniques developed on a flight support simulator. Spin modes at high and medium high angles of attack were identified, and recovery techniques were investigated. A flight support simulation of the airplane model mechanized with low speed wind tunnel data over an angle of attack range of + or - 90 deg. and an angle of sideslip range of + or - 40 deg. provided insight into the effects of altitude, stability, aerodynamic damping, and the operation of the augmented flight control system on spins. Aerodynamic derivatives determined from flight maneuvers were used to correlate model controllability with two proposed departure/spin design criteria.

Development and test results of a flight management algorithm for fuel conservative descents in a time-based metered traffic environment

NASA Technical Reports Server (NTRS)

Knox, C. E.; Cannon, D. G.

1980-01-01

A simple flight management descent algorithm designed to improve the accuracy of delivering an airplane in a fuel-conservative manner to a metering fix at a time designated by air traffic control was developed and flight tested. This algorithm provides a three dimensional path with terminal area time constraints (four dimensional) for an airplane to make an idle thrust, clean configured (landing gear up, flaps zero, and speed brakes retracted) descent to arrive at the metering fix at a predetermined time, altitude, and airspeed. The descent path was calculated for a constant Mach/airspeed schedule from linear approximations of airplane performance with considerations given for gross weight, wind, and nonstandard pressure and temperature effects. The flight management descent algorithm is described. The results of the flight tests flown with the Terminal Configured Vehicle airplane are presented.

Autonomous Flight Safety System

NASA Technical Reports Server (NTRS)

Simpson, James

2010-01-01

The Autonomous Flight Safety System (AFSS) is an independent self-contained subsystem mounted onboard a launch vehicle. AFSS has been developed by and is owned by the US Government. Autonomously makes flight termination/destruct decisions using configurable software-based rules implemented on redundant flight processors using data from redundant GPS/IMU navigation sensors. AFSS implements rules determined by the appropriate Range Safety officials.

First Spacelab flight - A status report of the joint ESA/NASA mission

NASA Technical Reports Server (NTRS)

Craft, H. G., Jr.; Sanfourche, J.-P.

1978-01-01

A general overview of the first Spacelab flight is presented and a table is given listing the payload composition. An accommodation study is presented with emphasis on the configuration, mass status, timeline, and experiment interface specifications. Also considered are flight and ground operations, safety factors, and payload specialists training for the first flight.

Development of Autonomous Optimal Cooperative Control in Relay Rover Configured Small Unmanned Aerial Systems

DTIC Science & Technology

2013-03-01

Unmanned Aircraft Systems Flight Plan that identified small unmanned aerial systems ( SUAS ) as “a profound technological...advances in small unmanned aerial systems ( SUAS ) cooperative control. The end state objective of the research effort was to flight test an autonomous...requirements were captured in the Unmanned Aircraft Systems Flight Plan . The flight plan

A LTA flight research vehicle. [technology assessment, airships

NASA Technical Reports Server (NTRS)

Nebiker, F. R.

1975-01-01

An Airship Flight Research Program is proposed. Major program objectives are summarized and a Modernized Navy ZPG3W Airship recommended as the flight test vehicle. The origin of the current interest in modern airship vehicles is briefly discussed and the major benefits resulting from the flight research program described. Airship configurations and specifications are included.

14 CFR 23.251 - Vibration and buffeting.

Code of Federal Regulations, 2012 CFR

2012-01-01

... interfere with the satisfactory control of the airplane or cause excessive fatigue to the flight crew. Stall... flight condition, including configuration changes during cruise, severe enough to interfere with the satisfactory control of the airplane or cause excessive fatigue to the flight crew. Stall warning buffeting...

Pilot-in-the-Loop Analysis of Propulsive-Only Flight Control Systems

NASA Technical Reports Server (NTRS)

Chou, Hwei-Lan; Biezad, Daniel J.

1996-01-01

Longitudinal control system architectures are presented which directly couple flight stick motions to throttle commands for a multi-engine aircraft. This coupling enables positive attitude control with complete failure of the flight control system. The architectures chosen vary from simple feedback gains to classical lead-lag compensators with and without prefilters. Each architecture is reviewed for its appropriateness for piloted flight. The control systems are then analyzed with pilot-in-the-loop metrics related to bandwidth required for landing. Results indicate that current and proposed bandwidth requirements should be modified for throttles only flight control. Pilot ratings consistently showed better ratings than predicted by analysis. Recommendations are made for more robust design and implementation. The use of Quantitative Feedback Theory for compensator design is discussed. Although simple and effective augmented control can be achieved in a wide variety of failed configurations, a few configuration characteristics are dominant for pilot-in-the-loop control. These characteristics will be tested in a simulator study involving failed flight controls for a multi-engine aircraft.

Flight-determined derivatives and dynamic characteristics for the HL-10 lifting body vehicle at subsonic and transonic Mach numbers

NASA Technical Reports Server (NTRS)

Strutz, L. W.

1972-01-01

The HL-10 lifting body stability and control derivatives were determined by using an analog-matching technique and compared with derivatives obtained from wind-tunnel results. The flight derivatives were determined as a function of angle of attack for a subsonic configuration at Mach 0.7 and for a transonic configuration at Mach 0.7, 0.9, and 1.2. At an angle of attack of 14 deg, data were obtained for a Mach number range from 0.6 to 1.4. The flight and wind-tunnel derivatives were in general agreement, with the possible exception of the longitudinal and lateral damping derivatives. Some differences were noted between the vehicle dynamic response characteristics calculated from flight-determined derivatives and those predicted by the wind-tunnel results. However, the only difference the pilots noted between the response of the vehicle in flight and the response of a simulator programed with wind-tunnel-predicted data was that the damping generally was higher in the flight vehicle.

Space Shuttle Ascent Flight Design Process: Evolution and Lessons Learned

NASA Technical Reports Server (NTRS)

Picka, Bret A.; Glenn, Christopher B.

2011-01-01

The Space Shuttle Ascent Flight Design team is responsible for defining a launch to orbit trajectory profile that satisfies all programmatic mission objectives and defines the ground and onboard reconfiguration requirements for this high-speed and demanding flight phase. This design, verification and reconfiguration process ensures that all applicable mission scenarios are enveloped within integrated vehicle and spacecraft certification constraints and criteria, and includes the design of the nominal ascent profile and trajectory profiles for both uphill and ground-to-ground aborts. The team also develops a wide array of associated training, avionics flight software verification, onboard crew and operations facility products. These key ground and onboard products provide the ultimate users and operators the necessary insight and situational awareness for trajectory dynamics, performance and event sequences, abort mode boundaries and moding, flight performance and impact predictions for launch vehicle stages for use in range safety, and flight software performance. These products also provide the necessary insight to or reconfiguration of communications and tracking systems, launch collision avoidance requirements, and day of launch crew targeting and onboard guidance, navigation and flight control updates that incorporate the final vehicle configuration and environment conditions for the mission. Over the course of the Space Shuttle Program, ascent trajectory design and mission planning has evolved in order to improve program flexibility and reduce cost, while maintaining outstanding data quality. Along the way, the team has implemented innovative solutions and technologies in order to overcome significant challenges. A number of these solutions may have applicability to future human spaceflight programs.

Altitude Performance Characteristics of Turbojet-engine Tail-pipe Burner with Variable-area Exhaust Nozzle Using Several Fuel Systems and Flame Holders

NASA Technical Reports Server (NTRS)

Johnson, Lavern A; Meyer, Carl L

1950-01-01

A tail-pipe burner with a variable-area exhaust nozzle was investigated. From five configurations a fuel-distribution system and a flame holder were selected. The best configuration was investigated over a range of altitudes and flight Mach numbers. For the best configuration, an increase in altitude lowered the augmented thrust ratio, exhaust-gas total temperature, and tail-pipe combustion efficiency, and raised the specific fuel consumption. An increase in flight Mach number raised the augmented thrust ratio but had no apparent effect on exhaust-gas total temperature, tail-pipe combustion efficiency, or specific fuel consumption.

Flight test results for several light, canard-configured airplanes

NASA Technical Reports Server (NTRS)

Brown, Philip W.

1987-01-01

Brief flight evaluations of two different, light, composite constructed, canard and winglet configured airplanes were performed to assess their handling qualities; one airplane was a single engine, pusher design and the other a twin engine, push-pull configuration. An emphasis was placed on the slow speed/high angle of attack region for both airplanes and on the engine-out regime for the twin. Mission suitability assessment included cockpit and control layout, ground and airborne handling qualities, and turbulence response. Very limited performance data was taken. Stall/spin tests and the effects of laminar flow loss on performance and handling qualities were assessed on an extended range, single engine pusher design.

Air traffic management evaluation tool

NASA Technical Reports Server (NTRS)

Sridhar, Banavar (Inventor); Chatterji, Gano Broto (Inventor); Schipper, John F. (Inventor); Bilimoria, Karl D. (Inventor); Grabbe, Shon (Inventor); Sheth, Kapil S. (Inventor)

2012-01-01

Methods for evaluating and implementing air traffic management tools and approaches for managing and avoiding an air traffic incident before the incident occurs. A first system receives parameters for flight plan configurations (e.g., initial fuel carried, flight route, flight route segments followed, flight altitude for a given flight route segment, aircraft velocity for each flight route segment, flight route ascent rate, flight route descent route, flight departure site, flight departure time, flight arrival time, flight destination site and/or alternate flight destination site), flight plan schedule, expected weather along each flight route segment, aircraft specifics, airspace (altitude) bounds for each flight route segment, navigational aids available. The invention provides flight plan routing and direct routing or wind optimal routing, using great circle navigation and spherical Earth geometry. The invention provides for aircraft dynamics effects, such as wind effects at each altitude, altitude changes, airspeed changes and aircraft turns to provide predictions of aircraft trajectory (and, optionally, aircraft fuel use). A second system provides several aviation applications using the first system. Several classes of potential incidents are analyzed and averted, by appropriate change en route of one or more parameters in the flight plan configuration, as provided by a conflict detection and resolution module and/or traffic flow management modules. These applications include conflict detection and resolution, miles-in trail or minutes-in-trail aircraft separation, flight arrival management, flight re-routing, weather prediction and analysis and interpolation of weather variables based upon sparse measurements. The invention combines these features to provide an aircraft monitoring system and an aircraft user system that interact and negotiate changes with each other.

Flight-Test Evaluation of the Longitudinal Stability and Control Characteristics of 0.5-Scale Models of the Fairchild Lark Pilotless-Aircraft Configuration: Standard Configuration with Wing Flaps Deflected 60 Degrees and Model having Tail in Line with Wings, TED No. NACA 2387

NASA Technical Reports Server (NTRS)

Stone, David G.

1947-01-01

Flight tests were conducted at the Flight Test Station of the Pilotless Aircraft Research Division at Wallop Island, Va., to determine the longitudinal control and stability characteristics of 0.5-scale models of the Fairchild Lark pilotless aircraft with the tail in line with the wings a d with the horizontal wing flaps deflected 60 deg. The data were obtained by the use of a telemeter and by radar tracking.

Flight test of ARINC 741 configuration low gain SATCOM system on Boeing 747-400 aircraft

NASA Technical Reports Server (NTRS)

Murphy, Timothy A.; Stapleton, Brian P.

1990-01-01

The Boeing company conducted a flight test of a SATCOM system similar to the ARINC 741 configuration on a production model 747-400. A flight plan was specifically designed to test the system over a wide variety of satellite elevations and aircraft attitudes as well as over land and sea. Interface bit errors, signal quality and aircraft position and navigational inputs were all recorded as a function of time. Special aircraft maneuvers were performed to demonstrate the potential for shadowing by aircraft structures. Both a compass rose test and the flight test indicated that shadowing from the tail is insignificant for the 747-400. However, satellite elevation angles below the aircraft horizon during banking maneuvers were shown to have a significant deleterious effect on SATCOM communications.

Flight test of ARINC 741 configuration low gain SATCOM system on Boeing 747-400 aircraft

NASA Astrophysics Data System (ADS)

Murphy, Timothy A.; Stapleton, Brian P.

The Boeing company conducted a flight test of a SATCOM system similar to the ARINC 741 configuration on a production model 747-400. A flight plan was specifically designed to test the system over a wide variety of satellite elevations and aircraft attitudes as well as over land and sea. Interface bit errors, signal quality and aircraft position and navigational inputs were all recorded as a function of time. Special aircraft maneuvers were performed to demonstrate the potential for shadowing by aircraft structures. Both a compass rose test and the flight test indicated that shadowing from the tail is insignificant for the 747-400. However, satellite elevation angles below the aircraft horizon during banking maneuvers were shown to have a significant deleterious effect on SATCOM communications.

Flight directors for STOl aircraft

NASA Technical Reports Server (NTRS)

Rabin, U. H.

1983-01-01

Flight director logic for flight path and airspeed control of a powered-lift STOL aircraft in the approach, transition, and landing configurations are developed. The methods for flight director design are investigated. The first method is based on the Optimal Control Model (OCM) of the pilot. The second method, proposed here, uses a fixed dynamic model of the pilot in a state space formulation similar to that of the OCM, and includes a pilot work-load metric. Several design examples are presented with various aircraft, sensor, and control configurations. These examples show the strong impact of throttle effectiveness on the performance and pilot work-load associated with manual control of powered-lift aircraft during approach. Improved performed and reduced pilot work-load can be achieved by using direct-lift-control to increase throttle effectiveness.

B-52B-008/DTV (Drop Test Vehicle) configuration 1 (with and without fins) flight test results - captive flight and drop test missions

NASA Technical Reports Server (NTRS)

Quade, D. A.

1978-01-01

The B-52B-008 drop test consisted of one takeoff roll to 60 KCAS, two captive flights to accomplish limited safety of flight flutter and structural demonstration testing, and seven drop test flights. Of the seven drop test missions, one flight was aborted due to the failure of the hook mechanism to release the drop test vehicle (DTV); but the other six flights successfully dropped the DTV.

Cryogenic fluid management program flight concept definition

NASA Technical Reports Server (NTRS)

Kroeger, Erich

1987-01-01

The Lewis Research Center's cryogenic fluid management program flight concept definition is presented in viewgraph form. Diagrams are given of the cryogenic fluid management subpallet and its configuration with the Delta launch vehicle. Information is given in outline form on feasibility studies, requirements definition, and flight experiments design.

Flapping response characteristics of hingeless rotor blades by a gereralized harmonic balance method

NASA Technical Reports Server (NTRS)

Peters, D. A.; Ormiston, R. A.

1975-01-01

Linearized equations of motion for the flapping response of flexible rotor blades in forward flight are derived in terms of generalized coordinates. The equations are solved using a matrix form of the method of linear harmonic balance, yielding response derivatives for each harmonic of the blade deformations and of the hub forces and moments. Numerical results and approximate closed-form expressions for rotor derivatives are used to illustrate the relationships between rotor parameters, modeling assumptions, and rotor response characteristics. Finally, basic hingeless rotor response derivatives are presented in tabular and graphical form for a wide range of configuration parameters and operating conditions.

Low-lift-to-drag-ratio approach and landing studies using a CV-990 airplane

NASA Technical Reports Server (NTRS)

Kock, B. M.; Fulton, F. L.; Drinkwater, F. J., III

1972-01-01

The results are presented of a flight-test program utilizing a CV-990 airplane, flow in low-lift-to-drag-ratio (L/D) configurations, to simulate terminal area operation, approach, and landing of large unpowered vehicles. The results indicate that unpowered approaches and landings are practical with vehicles of the size and performance characteristics of the proposed shuttle vehicle. Low L/D landings provided touchdown dispersion patterns acceptable for operation on runways of reasonable length. The dispersion pattern was reduced when guidance was used during the final approach. High levels of pilot proficiency were not required for acceptable performance.

Development of an integrated configuration management/flight director system for piloted STOL approaches

NASA Technical Reports Server (NTRS)

Hoh, R. H.; Klein, R. H.; Johnson, W. A.

1977-01-01

A system analysis method for the development of an integrated configuration management/flight director system for IFR STOL approaches is presented. Curved descending decelerating approach trajectories are considered. Considerable emphasis is placed on satisfying the pilot centered requirements (acceptable workload) as well as the usual guidance and control requirements (acceptable performance). The Augmentor Wing Jet STOL Research Aircraft was utilized to allow illustration by example, and to validate the analysis procedure via manned simulation.

Structural Design for a Neptune Aerocapture Mission

NASA Technical Reports Server (NTRS)

Dyke, R. Eric; Hrinda, Glenn A.

2004-01-01

A multi-center study was conducted in 2003 to assess the feasibility of and technology requirements for using aerocapture to insert a scientific platform into orbit around Neptune. The aerocapture technique offers a potential method of greatly reducing orbiter mass and thus total spacecraft launch mass by minimizing the required propulsion system mass. This study involved the collaborative efforts of personnel from Langley Research Center (LaRC), Johnson Space Flight Center (JSFC), Marshall Space Flight Center (MSFC), Ames Research Center (ARC), and the Jet Propulsion Laboratory (JPL). One aspect of this effort was the structural design of the full spacecraft configuration, including the ellipsled aerocapture orbiter and the in-space solar electric propulsion (SEP) module/cruise stage. This paper will discuss the functional and structural requirements for each of these components, some of the design trades leading to the final configuration, the loading environments, and the analysis methods used to ensure structural integrity. It will also highlight the design and structural challenges faced while trying to integrate all the mission requirements. Component sizes, materials, construction methods and analytical results, including masses and natural frequencies, will be presented, showing the feasibility of the resulting design for use in a Neptune aerocapture mission. Lastly, results of a post-study structural mass optimization effort on the ellipsled will be discussed, showing potential mass savings and their influence on structural strength and stiffness

Bootstrap Percolation on Homogeneous Trees Has 2 Phase Transitions

NASA Astrophysics Data System (ADS)

Fontes, L. R. G.; Schonmann, R. H.

2008-09-01

We study the threshold θ bootstrap percolation model on the homogeneous tree with degree b+1, 2≤ θ≤ b, and initial density p. It is known that there exists a nontrivial critical value for p, which we call p f , such that a) for p> p f , the final bootstrapped configuration is fully occupied for almost every initial configuration, and b) if p< p f , then for almost every initial configuration, the final bootstrapped configuration has density of occupied vertices less than 1. In this paper, we establish the existence of a distinct critical value for p, p c , such that 0< p c < p f , with the following properties: 1) if p≤ p c , then for almost every initial configuration there is no infinite cluster of occupied vertices in the final bootstrapped configuration; 2) if p> p c , then for almost every initial configuration there are infinite clusters of occupied vertices in the final bootstrapped configuration. Moreover, we show that 3) for p< p c , the distribution of the occupied cluster size in the final bootstrapped configuration has an exponential tail; 4) at p= p c , the expected occupied cluster size in the final bootstrapped configuration is infinite; 5) the probability of percolation of occupied vertices in the final bootstrapped configuration is continuous on [0, p f ] and analytic on ( p c , p f ), admitting an analytic continuation from the right at p c and, only in the case θ= b, also from the left at p f .

Experiments using electronic display information in the NASA terminal configured vehicle

NASA Technical Reports Server (NTRS)

Morello, S. A.

1980-01-01

The results of research experiments concerning pilot display information requirements and visualization techniques for electronic display systems are presented. Topics deal with display related piloting tasks in flight controls for approach-to-landing, flight management for the descent from cruise, and flight operational procedures considering the display of surrounding air traffic. Planned research of advanced integrated display formats for primary flight control throughout the various phases of flight is also discussed.

Thematic mapper flight model preshipment review data package. Volume 4: Appendix. Part B: Scan mirror assembly data

NASA Technical Reports Server (NTRS)

1982-01-01

Data from the thematic mapper scan mirror assembly (SMA) acceptance test are presented. Documentation includes: (1) a list of the acceptance test discrepancies; (2) flight 1 SMA test data book; (3) flight 1 SMA environmental report; (4) the configuration verification index; (5) the flight 1 SMA test failure reports; (6) the flight 1 data tapes log; and (7) the requests for deviation/waivers.

Comparing Methods for Dynamic Airspace Configuration

NASA Technical Reports Server (NTRS)

Zelinski, Shannon; Lai, Chok Fung

2011-01-01

This paper compares airspace design solutions for dynamically reconfiguring airspace in response to nominal daily traffic volume fluctuation. Airspace designs from seven algorithmic methods and a representation of current day operations in Kansas City Center were simulated with two times today's demand traffic. A three-configuration scenario was used to represent current day operations. Algorithms used projected unimpeded flight tracks to design initial 24-hour plans to switch between three configurations at predetermined reconfiguration times. At each reconfiguration time, algorithms used updated projected flight tracks to update the subsequent planned configurations. Compared to the baseline, most airspace design methods reduced delay and increased reconfiguration complexity, with similar traffic pattern complexity results. Design updates enabled several methods to as much as half the delay from their original designs. Freeform design methods reduced delay and increased reconfiguration complexity the most.

Static Longitudinal Stability of a Tandem-Coupled Bomber-Fighter Airplane Configuration Similar to One Proposed by Douglas Aircraft Company, Inc.

NASA Technical Reports Server (NTRS)

Hewes, Donald E.

1950-01-01

At the request of the Air Materiel Command, an investigation was made in the Langley free-flight tunnel to determine the longitudinal stability and control characteristics of models coupled together in a tandem configuration for aerial refueling similar to one proposed by the Douglas Aircraft Company, Inc. Static force tests were made with 1/20-scale models of the B-29 and F-80 airplanes to determine the effects of rigidly coupling the airplanes together. The Douglas configuration differs from the rigid configuration tested in that it provides for some freedom in pitch and vertical displacement. The force tests showed that, for the bomber alone, the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration with rigid coupling the aerodynamic center was 0.28 mean aerodynamic chord forward of the center of gravity of the combination (unstable). This reduction in stability was caused by the downwash of the bomber on the fighter. The pitching moment produced by elevator deflection of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel on models in a similar tandem configuration indicated that, with a hinged coupling permitting freedom in pitch, the stability of the combination was better than that obtained with a rigid coupling and was about the same as that for the bomber alone.

The development of an automated flight test management system for flight test planning and monitoring

NASA Technical Reports Server (NTRS)

Hewett, Marle D.; Tartt, David M.; Duke, Eugene L.; Antoniewicz, Robert F.; Brumbaugh, Randal W.

1988-01-01

The development of an automated flight test management system (ATMS) as a component of a rapid-prototyping flight research facility for AI-based flight systems concepts is described. The rapid-prototyping facility includes real-time high-fidelity simulators, numeric and symbolic processors, and high-performance research aircraft modified to accept commands for a ground-based remotely augmented vehicle facility. The flight system configuration of the ATMS includes three computers: the TI explorer LX and two GOULD SEL 32/27s.

Aircraft Configuration and Flight Crew Compliance with Procedures While Conducting Flight Deck Based Interval Management (FIM) Operations

NASA Technical Reports Server (NTRS)

Shay, Rick; Swieringa, Kurt A.; Baxley, Brian T.

2012-01-01

Flight deck based Interval Management (FIM) applications using ADS-B are being developed to improve both the safety and capacity of the National Airspace System (NAS). FIM is expected to improve the safety and efficiency of the NAS by giving pilots the technology and procedures to precisely achieve an interval behind the preceding aircraft by a specific point. Concurrently but independently, Optimized Profile Descents (OPD) are being developed to help reduce fuel consumption and noise, however, the range of speeds available when flying an OPD results in a decrease in the delivery precision of aircraft to the runway. This requires the addition of a spacing buffer between aircraft, reducing system throughput. FIM addresses this problem by providing pilots with speed guidance to achieve a precise interval behind another aircraft, even while flying optimized descents. The Interval Management with Spacing to Parallel Dependent Runways (IMSPiDR) human-in-the-loop experiment employed 24 commercial pilots to explore the use of FIM equipment to conduct spacing operations behind two aircraft arriving to parallel runways, while flying an OPD during high-density operations. This paper describes the impact of variations in pilot operations; in particular configuring the aircraft, their compliance with FIM operating procedures, and their response to changes of the FIM speed. An example of the displayed FIM speeds used incorrectly by a pilot is also discussed. Finally, this paper examines the relationship between achieving airline operational goals for individual aircraft and the need for ATC to deliver aircraft to the runway with greater precision. The results show that aircraft can fly an OPD and conduct FIM operations to dependent parallel runways, enabling operational goals to be achieved efficiently while maintaining system throughput.

NASA advanced aeronautics design solar powered remotely piloted vehicle

NASA Technical Reports Server (NTRS)

Elario, David S.; Guillmette, Neal H.; Lind, Gregory S.; Webster, Jonathan D.; Ferreira, Michael J.; Konstantakis, George C.; Marshall, David L.; Windt, Cari L.

1991-01-01

Environmental problems such as the depletion of the ozone layer and air pollution demand a change in traditional means of propulsion that is sensitive to the ecology. Solar powered propulsion is a favorable alternative that is both ecologically harmless as well as cost effective. Integration of solar energy into designs ranging from futuristic vehicles to heating is beneficial to society. The design and construction of a Multi-Purpose Remotely Piloted Vehicle (MPRPV) seeks to verify the feasibility of utilizing solar propulsion as a primary fuel source. This task has been a year long effort by a group of ten students, divided into five teams, each dealing with different aspects of the design. The aircraft was designed to take-off, climb to the design altitude, fly in a sustained figure-eight flight path, and cruise for approximately one hour. This mission requires flight at Reynolds numbers between 150,000 and 200,000 and demands special considerations in the aerodynamic design in order to achieve flight in this regime. Optimal performance requires a light weight configuration with both structural integrity and maximum power availability. The structure design and choice of solar cells for the propulsion was governed by the weight, efficiency, and cost considerations. The final design is a MPRPV weighting 35 N which cruises 7 m/s at the design altitude of 50 m. The configuration includes a wing composed of balsa and foam NACA 6409 airfoil sections and carbon fiber spars, a tail of similar construction, and a truss structure fuselage. The propulsion system consists of 98 10 percent efficient solar cells donated by Mobil Solar, a NiCad battery for energy storage, and a folding propeller regulated by a lightweight and efficient control system. The airfoils and propeller chosen for the design were research and tested during the design process.

Solar powered multipurpose remotely powered aircraft

NASA Technical Reports Server (NTRS)

1991-01-01

Environmental problems such as the depletion of the ozone layer and air pollution demand a change in traditional means of propulsion that is sensitive to the ecology. Solar-powered propulsion is a favorable alternative that is both ecologically harmless as well as cost effective. Integration of solar energy into designs ranging from futuristic vehicles to heating is, therefore, beneficial to society. The design and construction of a Multipurpose Remotely Piloted Vehicle (MPRPV) seeks to verify the feasibility of using solar propulsion as a primary fuel source. This task has been a year-long effort by a group of eight students, divided into four teams, each dealing with different aspects of the design. The aircraft was designed to take off, climb to the design altitude, fly in a sustained figure-eight flight path, and cruise for approximately one hour. This mission requires flight at Reynolds numbers between 150,000 and 200,000 and demands special considerations in the aerodynamic design to achieve flight in this regime. Optimal performance requires a lightweight configuration with both structural integrity and maximum power availability. The structural design and choice of solar cells for the propulsion were governed by weight, efficiency, and cost considerations. The final design is an MPRPV weighing 35 N that cruises at 7 m/s at the design altitude of 50 m. The configuration includes a wing composed of balsa and foam NACA 6409 airfoil sections and carbon fiber spars, a tail of similar construction, and a truss structure fuselage. The propulsion system consists of 98 12.5 percent-efficient solar cells donated by Mobil Solar, a NiCad battery for energy storage, and a folding propeller regulated by a lightweight and efficient control system. The airfoils and propeller chosen for the design were researched and tested during the design process.

Multi-reactor power system configurations for multimegawatt nuclear electric propulsion

NASA Technical Reports Server (NTRS)

George, Jeffrey A.

1991-01-01

A modular, multi-reactor power system and vehicle configuration for piloted nuclear electric propulsion (NEP) missions to Mars is presented. Such a design could provide enhanced system and mission reliability, allowing a comfortable safety margin for early manned flights, and would allow a range of piloted and cargo missions to be performed with a single power system design. Early use of common power modules for cargo missions would also provide progressive flight experience and validation of standardized systems for use in later piloted applications. System and mission analysis are presented to compare single and multi-reactor configurations for piloted Mars missions. A conceptual design for the Hydra modular multi-reactor NEP vehicle is presented.

Scramjet nozzle design and analysis as applied to a highly integrated hypersonic research airplane

NASA Technical Reports Server (NTRS)

Small, W. J.; Weidner, J. P.; Johnston, P. J.

1974-01-01

The configuration and performance of the propulsion system for the hypersonic research vehicle are discussed. A study of the interactions between propulsion and aerodynamics of the highly integrated vehicle was conducted. The hypersonic research vehicle is configured to test the technology of structural and thermal protection systems concepts and the operation of the propulsion system under true flight conditions for most of the hypersonic flight regime. The subjects considered are: (1) research vehicle and scramjet engine configurations to determine fundamental engine sizing constraints, (2) analytical methods for computing airframe and propulsion system components, and (3) characteristics of a candidate nozzle to investigate vehicle stability and acceleration performance.

SOFIA Closed- and Open-Door Aerodynamic Analyses

NASA Technical Reports Server (NTRS)

Cumming, Stephen; Frederick, Mike; Smith, Mark

2012-01-01

Work to evaluate the aerodynamic characteristics and the cavity acoustic environment of the SOFIA (Stratospheric Observatory for Infrared Astronomy) airplane has been completed. The airplane has been evaluated in its closed-door configuration, as well as several open-door configurations. Work performed included: acoustic analysis tool development, cavity acoustic evaluation, stability and control parameter estimation, air data calibration, and external flow evaluation. Qualitative airflow data were obtained during the closed- and open-door flights using tufts on the aft portion of the fuselage. Video was taken from a chase plane. This video was analyzed for various flight conditions, and general flow descriptions of the aft fuselage of the 747SP were developed for the different closed and open door configurations.

Virtual Flight Demonstration of the Stratospheric Dual-Aircraft Platform

NASA Technical Reports Server (NTRS)

Engblom, W. A.; Decker, R. K.

2016-01-01

A baseline configuration for the dual-aircraft platform (DAP) concept is described and evaluated in a physics-based flight dynamics simulations for two month-long missions as a communications relay in the lower stratosphere above central Florida. The DAP features two unmanned aerial vehicles connected via a long adjustable cable which effectively sail back-and-forth using wind velocity gradients and solar energy. Detailed atmospheric profiles in the vicinity of 60,000-ft derived from archived data measured by the 50-Mhz Doppler Radar Wind Profiler at Cape Canaveral are used in the flight simulations. An overview of the novel guidance and flight control strategies are provided. The energy-usage of the baseline configuration during month-long stationkeeping missions (i.e., within 150-mile radius of downtown Orlando) is characterized and compared to that of a pure solar aircraft.

Multidimensional Effects on Ignition, Transition, and Flame Spread in Microgravity

NASA Technical Reports Server (NTRS)

Kashiwagi, T.; Mell, W. E.; Nakamura, Y.; Olson, S. L.; Baum, H. R.; McGrattan, K. B.

2001-01-01

Localized ignition is initiated by an external radiant source at the middle of a thermally thin sample under external slow flow, simulating fire initiation in a spacecraft with a slow ventilation flow. Two ignition configurations are simulated, one across the sample surface creating a line shaped flame front (two-dimensional, 2-D, configuration) and the other a small circular ignition (three-dimensional, 3-D, configuration). Ignition, subsequent transition to simultaneously upstream and downstream flame spread, and flame growth behavior are studied experimentally and theoretically. Details of our theoretical models and numerical techniques can be found in previous publications. The effects of the sample width on the transition and subsequent flame spread, and flame spread along open edges of a thermally thin paper sample are determined. Experimental observations of flame spread phenomena were conducted in the 10 s drop tower and also on the space shuttle STS-75 flight to determine the effects of oxygen concentration and external flow velocity on flame spread rate and flame growth pattern. Finally, effects of confinement in a small test chamber on the transition and subsequent flame spread are examined. The results of these studies are briefly reported.

A candidate V/STOL research aircraft design concept using an S-3A aircraft and 2 Pegasus 11 engines

NASA Technical Reports Server (NTRS)

Lampkin, B. A.

1980-01-01

A candidate V/STOL research aircraft concept which uses an S-3A airframe and two Pegasus 11 engines was studied to identify a feasible V/STOL national flight facility that could be obtained at the lowest possible cost for the demonstration of V/STOL technology, inflight simulation, and flight research. The rationale for choosing the configuration, a description of the configuration, and the capability of a fully developed aircraft are discussed.

Flight Determination of the Longitudinal Stability in Accelerated Maneuvers at Transonic Speeds for the Douglas D-558-II Research Airplane Including the Effects of an Outboard Wing Fence

NASA Technical Reports Server (NTRS)

Fischel, Jack; Nugent, Jack

1953-01-01

The results of transonic flight measurements of the longitudinal stability characteristics of the Douglas D-558-II research airplane in the original configuration and with outboard fences mounted on the wings are presented. The levels of normal-force coefficient at which the stability decreases and pitch-up starts have been determined for both airplane configurations at Mach numbers up to about 0.94.

Small UAV Automatic Ground Collision Avoidance System Design Considerations and Flight Test Results

NASA Technical Reports Server (NTRS)

Sorokowski, Paul; Skoog, Mark; Burrows, Scott; Thomas, SaraKatie

2015-01-01

The National Aeronautics and Space Administration (NASA) Armstrong Flight Research Center Small Unmanned Aerial Vehicle (SUAV) Automatic Ground Collision Avoidance System (Auto GCAS) project demonstrated several important collision avoidance technologies. First, the SUAV Auto GCAS design included capabilities to take advantage of terrain avoidance maneuvers flying turns to either side as well as straight over terrain. Second, the design also included innovative digital elevation model (DEM) scanning methods. The combination of multi-trajectory options and new scanning methods demonstrated the ability to reduce the nuisance potential of the SUAV while maintaining robust terrain avoidance. Third, the Auto GCAS algorithms were hosted on the processor inside a smartphone, providing a lightweight hardware configuration for use in either the ground control station or on board the test aircraft. Finally, compression of DEM data for the entire Earth and successful hosting of that data on the smartphone was demonstrated. The SUAV Auto GCAS project demonstrated that together these methods and technologies have the potential to dramatically reduce the number of controlled flight into terrain mishaps across a wide range of aviation platforms with similar capabilities including UAVs, general aviation aircraft, helicopters, and model aircraft.

Flight Results of the Chandra X-ray Observatory Inertial Upper Stage Space Mission

NASA Technical Reports Server (NTRS)

Tillotson, R.; Walter, R.

2000-01-01

Under contract to NASA, a specially configured version of the Boeing developed Inertial Upper Stage (IUS) booster was provided by Boeing to deliver NASA's 1.5 billion dollar Chandra X-Ray Observatory satellite into a highly elliptical transfer orbit from a Shuttle provided circular park orbit. Subsequently, the final orbit of the Chandra satellite was to be achieved using the Chandra Integral Propulsion System (IPS) through a series of IPS burns. On 23 July 1999 the Shuttle Columbia (STS-93) was launched with the IUS/Chandra stack in the Shuttle payload bay. Unfortunately, the Shuttle Orbiter was unexpectantly inserted into an off-nominal park orbit due to a Shuttle propulsion anomaly occurring during ascent. Following the IUS/Chandra on-orbit deployment from the Shuttle, at seven hours from liftoff, the flight proven IUS GN&C system successfully injected Chandra into the targeted transfer orbit, in spite of the off-nominal park orbit. This paper describes the IUS GN&C system, discusses the specific IUS GN&C mission data load development, analyses and testing for the Chandra mission, and concludes with a summary of flight results for the IUS part of the Chandra mission.

Digital flight control research

NASA Technical Reports Server (NTRS)

Potter, J. E.; Stern, R. G.; Smith, T. B.; Sinha, P.

1974-01-01

The results of studies which were undertaken to contribute to the design of digital flight control systems, particularly for transport aircraft are presented. In addition to the overall design considerations for a digital flight control system, the following topics are discussed in detail: (1) aircraft attitude reference system design, (2) the digital computer configuration, (3) the design of a typical digital autopilot for transport aircraft, and (4) a hybrid flight simulator.

Water tunnel flow visualization study of a 4.4 percent scale X-31 forebody

NASA Technical Reports Server (NTRS)

Cobleigh, Brent R.; Delfrate, John

1994-01-01

A water-tunnel test of a 4.4 percent-scale, forebody-only model of the X-31 aircraft with different forebody strakes and nosebooms has been performed in the Flow Visualization Facility at the NASA Dryden Flight Research Center. The focus of the study was to determine the relative effects of the different configurations on the stability and symmetry of the high-angle-of-attack forebody vortex flow field. The clean, noseboom-off configuration resisted the development of asymmetries in the primary vortices through 70 deg angle of attack. The wake of the X-31 flight test noseboom configuration significantly degraded the steadiness of the primary vortex cores and promoted asymmetries. An alternate L-shaped noseboom mounted underneath the forebody had results similar to those seen with the configuration, enabling stable, symmetrical vortices up to 70 deg angle of attack. The addition of strakes near the radome tip along the waterline increased the primary vortex strength while it simultaneously caused the vortex breakdown location co move forward. Forebody strakes did not appear to significantly reduce the asymmetries in the forebody vortex field in the presence of the flight test noseboom.

A Limited In-Flight Evaluation of the Constant Current Loop Strain Measurement Method

NASA Technical Reports Server (NTRS)

Olney, Candida D.; Collura, Joseph V.

1997-01-01

For many years, the Wheatstone bridge has been used successfully to measure electrical resistance and changes in that resistance. However, the inherent problem of varying lead wire resistance can cause errors when the Wheatstone bridge is used to measure strain in a flight environment. The constant current loop signal-conditioning card was developed to overcome that difficulty. This paper describes a limited evaluation of the constant current loop strain measurement method as used in the F-16XL ship 2 Supersonic Laminar Flow Control flight project. Several identical strain gages were installed in close proximity on a shock fence which was mounted under the left wing of the F- 1 6XL ship 2. Two strain gage bridges were configured using the constant current loop, and two were configured using the Wheatstone bridge circuitry. Flight data comparing the output from the constant current loop configured gages to that of the Wheatstone bridges with respect to signal output, error, and noise are given. Results indicate that the constant current loop strain measurement method enables an increased output, unaffected by lead wire resistance variations, to be obtained from strain gages.

Automated Tetrahedral Mesh Generation for CFD Analysis of Aircraft in Conceptual Design

NASA Technical Reports Server (NTRS)

Ordaz, Irian; Li, Wu; Campbell, Richard L.

2014-01-01

The paper introduces an automation process of generating a tetrahedral mesh for computational fluid dynamics (CFD) analysis of aircraft configurations in early conceptual design. The method was developed for CFD-based sonic boom analysis of supersonic configurations, but can be applied to aerodynamic analysis of aircraft configurations in any flight regime.

Experimental/Analytical Characterization of the RBCC Rocket-Ejector Mode

NASA Technical Reports Server (NTRS)

Ruf, J. H.; Lehman, M.; Pal, S.; Santoro, R. J.

2000-01-01

The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations are being conducted at Marshall Space Flight Center to benchmark the FDNS code for RBCC engine operations for such configurations. The primary fluid physics of interests are the mixing and interaction of the rocket plume and secondary flow, subsequent combustion of the fuel rich rocket exhaust with the secondary flow and combustion of the injected afterburner flow. The CFD results are compared to static pressure along the RBCC duct walls, Raman Spectroscopy specie distribution data at several axial locations, net engine thrust and entrained air for the SLS cases. The CFD results compare reasonably well with the experimental results.

Product assurance policies and procedures for flight dynamics software development

NASA Technical Reports Server (NTRS)

Perry, Sandra; Jordan, Leon; Decker, William; Page, Gerald; Mcgarry, Frank E.; Valett, Jon

1987-01-01

The product assurance policies and procedures necessary to support flight dynamics software development projects for Goddard Space Flight Center are presented. The quality assurance and configuration management methods and tools for each phase of the software development life cycles are described, from requirements analysis through acceptance testing; maintenance and operation are not addressed.

14 CFR 135.171 - Shoulder harness installation at flight crewmember stations.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Shoulder harness installation at flight... ON BOARD SUCH AIRCRAFT Aircraft and Equipment § 135.171 Shoulder harness installation at flight... configuration, excluding any pilot seat, of 10 seats or more unless it is equipped with an approved shoulder...

14 CFR 135.171 - Shoulder harness installation at flight crewmember stations.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Shoulder harness installation at flight... ON BOARD SUCH AIRCRAFT Aircraft and Equipment § 135.171 Shoulder harness installation at flight... configuration, excluding any pilot seat, of 10 seats or more unless it is equipped with an approved shoulder...

14 CFR 135.171 - Shoulder harness installation at flight crewmember stations.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Shoulder harness installation at flight... ON BOARD SUCH AIRCRAFT Aircraft and Equipment § 135.171 Shoulder harness installation at flight... configuration, excluding any pilot seat, of 10 seats or more unless it is equipped with an approved shoulder...

14 CFR 135.171 - Shoulder harness installation at flight crewmember stations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Shoulder harness installation at flight... ON BOARD SUCH AIRCRAFT Aircraft and Equipment § 135.171 Shoulder harness installation at flight... configuration, excluding any pilot seat, of 10 seats or more unless it is equipped with an approved shoulder...

14 CFR 135.171 - Shoulder harness installation at flight crewmember stations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Shoulder harness installation at flight... ON BOARD SUCH AIRCRAFT Aircraft and Equipment § 135.171 Shoulder harness installation at flight... configuration, excluding any pilot seat, of 10 seats or more unless it is equipped with an approved shoulder...

Propulsion system tests on a full scale Centaur vehicle to investigate 3-burn mission capability of the D-lT configuration

NASA Technical Reports Server (NTRS)

Groesbeck, W. A.; Baud, K. M.; Lacovic, R. F.; Tabata, W. K.; Szabo, S. V., Jr.

1974-01-01

Propulsion system tests were conducted on a full scale Centaur vehicle to investigate system capability of the proposed D-lT configuration for a three-burn mission. This particular mission profile requires that the engines be capable of restarting and firing for a final maneuver after a 5-1/2-hour coast to synchronous orbit. The thermal conditioning requirements of the engine and propellant feed system components for engine start under these conditions were investigated. Performance data were also obtained on the D-lT type computer controlled propellant tank pressurization system. The test results demonstrated that the RL-10 engines on the Centaur vehicle could be started and run reliably after being thermally conditioned to predicted engine start conditions for a one, two and three burn mission. Investigation of the thermal margins also indicated that engine starts could be accomplished at the maximum predicted component temperature conditions with prestart durations less than planned for flight.

A Common DPU Platform for ESA JUICE Mission Instruments

NASA Astrophysics Data System (ADS)

Aberg, Martin; Hellstrom, Daniel; Samuelsson, Arne; Torelli, Felice

2016-08-01

This paper describes the resulting hardware and software platform based on GR712RC [1] LEON3-FT that Cobham Gaisler developed in accordance with the common system requirements of the ten scientific instruments on-board the ESA JUICE spacecraft destined the Jupiter system [8].The radiation hardened DPU platform features EDAC protected boot, application memory and working memory of configurable sizes and SpaceWire, FPGA I/O-32/16/8, GPIO, UART and SPI I/O interfaces. The design has undergone PSA, Risk, WCA, Radiation analyses etc. to justify component and design choices resulting in a robust design that can be used in spacecrafts requiring a total dose up to 100krad(Si). The prototype board manufactured uses engineering models of the flight components to ensure that development is representative.Validated boot, standby and driver software accommodates the various DPU platform configurations. The boot performs low-level DPU initialization, standby handles OBC SpaceWire communication and finally the loading and executing of application images typically stored in the non-volatile application memory.

Experimental investigation of a 0.15 scale model of a conformal variable-ramp inlet for the F-16 airplane

NASA Technical Reports Server (NTRS)

Hawkins, J. E.

1980-01-01

A 0.15 scale model of a proposed conformal variable-ramp inlet for the Multirole Fighter was tested from Mach 0.8 to 2.2 at a wide range of angles of attack and sideslip. Inlet ramp angle was varied to optimize ramp angle as a function of engine airflow, Mach number, angle of attack, and angle of sideslip. Several inlet configuration options were investigated to study their effects on inlet operation and to establish the final flight configuration. These variations were cowl sidewall cutback, cowl lip bluntness, boundary layer bleed, and first-ramp leading edge shape. Diagnostic and engine face instrumentation were used to evaluate inlet operation at various inlet stations and at the inlet/engine interface. Pressure recovery and stability of the inlet were satisfactory for the proposed application. On the basis of an engine stability audit of the worst-case instantaneous distortion patterns, no inlet/engine compatibility problems are expected for normal operations.

Flight effects on noise generated by the JT8D-17 engine in a quiet nacelle and a conventional nacelle as measured in the NASA-Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Strout, F. G.

1976-01-01

A JT8D-17 turbofan engine was tested in the NASA-Ames 40- by 80-foot wind tunnel to determine flight effects on jet and fan noise. Baseline, quiet nacelle with 20-lobe ejector/suppressor, and internal mixer configurations were tested over a range of engine power settings and tunnel velocities. Flight effects derived from the 40- by 80-foot wind tunnel test are compared with 727/JT8D flight test data and with model data obtained in a smaller wind tunnel. Procedures are defined for measuring noise data in a wind tunnel relatively near the sources and analyzing the results to obtain far-field flight effects. Wind tunnel and 727 flight test noise results compare favorably for both the baseline and quiet nacelle configurations. Two reports are provided, including a comprehensive version with extensive test results and analysis and the subject summary version that emphasizes data analysis and program finding.

Free-Flight-Tunnel Investigation of the Dynamic Stability and Control Characteristics of a Chance Vought F7U-3 Airplane in Towed Flight

NASA Technical Reports Server (NTRS)

Grana, David C.; Shanks, Robert E.

1952-01-01

As part of a program to determine the feasibility of using a fighter airplane as a parasite in combination with a Consolidated Vultee RB-36 for long-range reconnaissance missions (project FICON), an experimental investigation has been made in the Langley free-flight tunnel to determine the dynamic stability and control characteristics of a 1/17.5-scale model of a Chance Vought F7U-3 airplane in several tow configurations. The investigation consisted of flight tests in which the model was towed from a strut in the tunnel by a towline and by a direct coupling which provided complete angular freedom. The tests with the direct coupling also included a study of the effect of spring restraint in roll in order to simulate approximately the proposed full-scale arrangement in which the only freedom is that permitted by the flexibility of the launching and retrieving trapeze carried by the-bomber. For the tow configurations in which a towline was used (15 and 38 feet full scale), the model had a very unstable lateral oscillation which could not be controlled. The stability was also unsatisfactory for the tow configuration in Which the model was coupled directly to the strut with complete angular freedom. When spring restraint in roll was added, however, the stability was satisfactory. The use of the yaw damper which increased the damping in yaw to about six times the normal value of the model appeared to have no appreciable effect on the lateral oscillations in the towline configurations, but produced a slight improvement in the case of the direct coupling configurations. The longitudinal stability was satisfactory for those cases in which the lateral stability was good enough to permit study of longitudinal motions.

Control theory analysis of a three-axis VTOL flight director. M.S. Thesis - Pennsylvania State Univ.

NASA Technical Reports Server (NTRS)

Niessen, F. R.

1971-01-01

A control theory analysis of a VTOL flight director and the results of a fixed-based simulator evaluation of the flight-director commands are discussed. The VTOL configuration selected for this study is a helicopter-type VTOL which controls the direction of the thrust vector by means of vehicle-attitude changes and, furthermore, employs high-gain attitude stabilization. This configuration is the same as one which was simulated in actual instrument flight tests with a variable stability helicopter. Stability analyses are made for each of the flight-director commands, assuming a single input-output, multi-loop system model for each control axis. The analyses proceed from the inner-loops to the outer-loops, using an analytical pilot model selected on the basis of the innermost-loop dynamics. The time response of the analytical model of the system is primarily used to adjust system gains, while root locus plots are used to identify dominant modes and mode interactions.

Self-Contained Avionics Sensing and Flight Control System for Small Unmanned Aerial Vehicle

NASA Technical Reports Server (NTRS)

Ingham, John C. (Inventor); Shams, Qamar A. (Inventor); Logan, Michael J. (Inventor); Fox, Robert L. (Inventor); Fox, legal representative, Melanie L. (Inventor); Kuhn, III, Theodore R. (Inventor); Babel, III, Walter C. (Inventor); Fox, legal representative, Christopher L. (Inventor); Adams, James K. (Inventor); Laughter, Sean A. (Inventor)

2011-01-01

A self-contained avionics sensing and flight control system is provided for an unmanned aerial vehicle (UAV). The system includes sensors for sensing flight control parameters and surveillance parameters, and a Global Positioning System (GPS) receiver. Flight control parameters and location signals are processed to generate flight control signals. A Field Programmable Gate Array (FPGA) is configured to provide a look-up table storing sets of values with each set being associated with a servo mechanism mounted on the UAV and with each value in each set indicating a unique duty cycle for the servo mechanism associated therewith. Each value in each set is further indexed to a bit position indicative of a unique percentage of a maximum duty cycle for the servo mechanism associated therewith. The FPGA is further configured to provide a plurality of pulse width modulation (PWM) generators coupled to the look-up table. Each PWM generator is associated with and adapted to be coupled to one of the servo mechanisms.

Correlation study of theoretical and experimental results for spin tests of a 1/10 scale radio control model

NASA Technical Reports Server (NTRS)

Bihrle, W., Jr.

1976-01-01

A correlation study was conducted to determine the ability of current analytical spin prediction techniques to predict the flight motions of a current fighter airplane configuration during the spin entry, the developed spin, and the spin recovery motions. The airplane math model used aerodynamics measured on an exact replica of the flight test model using conventional static and forced-oscillation wind-tunnel test techniques and a recently developed rotation-balance test apparatus capable of measuring aerodynamics under steady spinning conditions. An attempt was made to predict the flight motions measured during stall/spin flight testing of an unpowered, radio-controlled model designed to be a 1/10 scale, dynamically-scaled model of a current fighter configuration. Comparison of the predicted and measured flight motions show that while the post-stall and spin entry motions were not well-predicted, the developed spinning motion (a steady flat spin) and the initial phases of the spin recovery motion are reasonably well predicted.

Visually guided control of movement in the context of multimodal stimulation

NASA Technical Reports Server (NTRS)

Riccio, Gary E.

1991-01-01

Flight simulation has been almost exclusively concerned with simulating the motions of the aircraft. Physically distinct subsystems are often combined to simulate the varieties of aircraft motion. Visual display systems simulate the motion of the aircraft relative to remote objects and surfaces (e.g., other aircraft and the terrain). 'Motion platform' simulators recreate aircraft motion relative to the gravitoinertial vector (i.e., correlated rotation and tilt as opposed to the 'coordinated turn' in flight). 'Control loaders' attempt to simulate the resistance of the aerodynamic medium to aircraft motion. However, there are few operational systems that attempt to simulate the motion of the pilot relative to the aircraft and the gravitoinertial vector. The design and use of all simulators is limited by poor understanding of postural control in the aircraft and its effect on the perception and control of flight. Analysis of the perception and control of flight (real or simulated) must consider that: (1) the pilot is not rigidly attached to the aircraft; and (2) the pilot actively monitors and adjusts body orientation and configuration in the aircraft. It is argued that this more complete approach to flight simulation requires that multimodal perception be considered as the rule rather than the exception. Moreover, the necessity of multimodal perception is revealed by emphasizing the complementarity rather than the redundancy among perceptual systems. Finally, an outline is presented for an experiment to be conducted at NASA ARC. The experiment explicitly considers possible consequences of coordination between postural and vehicular control.

Numerically simulated comparative performance of a scramjet and shcramjet at Mach 11

NASA Astrophysics Data System (ADS)

Chan, Jonathan

This study investigates the design and aeropropulsive performance of a complete, hydrogen powered, shock-induced combustion ramjet (shcramjet) at a flight Mach number of 11 and altitude of 34.5 km. The design includes a Prandtl-Meyer compression inlet, cantilevered ramp fuel injectors, a shock-inducing wedge and a divergent nozzle. Numerical studies are undertaken using the WARP code that solves the three-dimensional Favre-averaged Navier-Stokes equations closed by the Wilcox k-o turbulence model and the Jachimowski H2/air chemical kinetics model. Studies of fuel injection properties, mixing duct length, combustor wedge and nozzle geometry are completed to maximize the overall performance of the vehicle. The final shcramjet configuration generates a specific impulse of 1110 s. A comparison is undertaken with a scramjet vehicle at identical flight conditions and using many of the same components. The comparable scramjet generates a higher specific impulse of 1450 s although it is significantly larger and therefore heavier.

Effect of cavity configuration on kerosene spark ignition in a scramjet combustor at Ma 4.5 flight condition

NASA Astrophysics Data System (ADS)

Bao, Heng; Zhou, Jin; Pan, Yu

2015-12-01

Spark ignition experiments of liquid kerosene are conducted in a scramjet model equipped with dual-cavities at Ma 4.5 flight condition with a stagnation temperature of 1032 K. The ignition ability of two cavities with different length is compared and analyzed based on the wall pressure distribution along the combustor and the thrust evolution. The experimental results indicate that the longer cavity (L/D=7) is more suitable than the smaller cavity (L/D=5) in spark ignition. When employing the smaller cavity, three steady combustion states are observed after spark ignition. The concept of 'local flame' is adopted to explain the expanding problem of weak combustion. The local equivalence ratio in the shear layer is the dominated factor in determining the developing process of local flame. The final steady combustion mode of the combustor is dependent on the flame developing process. When employing the longer cavity, the establishment of intense combustion state can be much easier.

Aerothermodynamic testing requirements for future space transportation systems

NASA Technical Reports Server (NTRS)

Paulson, John W., Jr.; Miller, Charles G., III

1995-01-01

Aerothermodynamics, encompassing aerodynamics, aeroheating, and fluid dynamic and physical processes, is the genesis for the design and development of advanced space transportation vehicles. It provides crucial information to other disciplines involved in the development process such as structures, materials, propulsion, and avionics. Sources of aerothermodynamic information include ground-based facilities, computational fluid dynamic (CFD) and engineering computer codes, and flight experiments. Utilization of this triad is required to provide the optimum requirements while reducing undue design conservatism, risk, and cost. This paper discusses the role of ground-based facilities in the design of future space transportation system concepts. Testing methodology is addressed, including the iterative approach often required for the assessment and optimization of configurations from an aerothermodynamic perspective. The influence of vehicle shape and the transition from parametric studies for optimization to benchmark studies for final design and establishment of the flight data book is discussed. Future aerothermodynamic testing requirements including the need for new facilities are also presented.

Wind-tunnel investigation of the flight characteristics of a canard general-aviation airplane configuration

NASA Technical Reports Server (NTRS)

Satran, D. R.

1986-01-01

A 0.36-scale model of a canard general-aviation airplane with a single pusher propeller and winglets was tested in the Langley 30- by 60-Foot Wind Tunnel to determine the static and dynamic stability and control and free-flight behavior of the configuration. Model variables made testing of the model possible with the canard in high and low positions, with increased winglet area, with outboard wing leading-edge droop, with fuselage-mounted vertical fin and rudder, with enlarged rudders, with dual deflecting rudders, and with ailerons mounted closer to the wing tips. The basic model exhibited generally good longitudinal and lateral stability and control characteristics. The removal of an outboard leading-edge droop degraded roll damping and produced lightly damped roll (wing rock) oscillations. In general, the model exhibited very stable dihedral effect but weak directional stability. Rudder and aileron control power were sufficiently adequate for control of most flight conditions, but appeared to be relatively weak for maneuvering compared with those of more conventionally configured models.

Design and development of a quad copter (UMAASK) using CAD/CAM/CAE

NASA Astrophysics Data System (ADS)

Manarvi, Irfan Anjum; Aqib, Muhammad; Ajmal, Muhammad; Usman, Muhammad; Khurshid, Saqib; Sikandar, Usman

Micro flying vehicles1 (MFV) have become a popular area of research due to economy of production, flexibility of launch and variety of applications. A large number of techniques from pencil sketching to computer based software are being used for designing specific geometries and selection of materials to arrive at novel designs for specific requirements. Present research was focused on development of suitable design configuration using CAD/CAM/CAE tools and techniques. A number of designs were reviewed for this purpose. Finally, rotary wing Quadcopter flying vehicle design was considered appropriate for this research. Performance requirements were planned as approximately 10 meters ceiling, weight less than 500grams and ability to take videos and pictures. Parts were designed using Finite Element Analysis, manufactured using CNC machines and assembled to arrive at final design named as UMAASK. Flight tests were carried out which confirmed the design requirements.

Laminar flow control SPF/08 feasibility demonstration

NASA Astrophysics Data System (ADS)

Ecklund, R. C.; Williams, N. R.

1981-10-01

The feasibility of applying superplastic forming/diffusion bonding (SPF/DB) technology to laminar flow control (LFC) system concepts was demonstrated. Procedures were developed to produce smooth, flat titanium panels, using thin -0.016 inch sheets, meeting LFC surface smoothness requirements. Two large panels 28 x 28 inches were fabricated as final demonstration articles. The first was flat on the top and bottom sides demonstrating the capability of the tooling and the forming and diffusion bonding procedures to produce flat, defect free surfaces. The second panel was configurated for LFC porous panel treatment by forming channels with dimpled projections on the top side. The projections were machined away leaving holes extending into the panel. A perforated titanium sheet was adhesively bonded over this surface to complete the LFC demonstration panel. The final surface was considered flat enough to meet LFC requirements for a jet transport aircraft in cruising flight.

Autonomous Flight Safety System - Phase III

NASA Technical Reports Server (NTRS)

2008-01-01

The Autonomous Flight Safety System (AFSS) is a joint KSC and Wallops Flight Facility project that uses tracking and attitude data from onboard Global Positioning System (GPS) and inertial measurement unit (IMU) sensors and configurable rule-based algorithms to make flight termination decisions. AFSS objectives are to increase launch capabilities by permitting launches from locations without range safety infrastructure, reduce costs by eliminating some downrange tracking and communication assets, and reduce the reaction time for flight termination decisions.

Aeroservoelastic Modeling and Validation of a Thrust-Vectoring F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Brenner, Martin J.

1996-01-01

An F/A-18 aircraft was modified to perform flight research at high angles of attack (AOA) using thrust vectoring and advanced control law concepts for agility and performance enhancement and to provide a testbed for the computational fluid dynamics community. Aeroservoelastic (ASE) characteristics had changed considerably from the baseline F/A-18 aircraft because of structural and flight control system amendments, so analyses and flight tests were performed to verify structural stability at high AOA. Detailed actuator models that consider the physical, electrical, and mechanical elements of actuation and its installation on the airframe were employed in the analysis to accurately model the coupled dynamics of the airframe, actuators, and control surfaces. This report describes the ASE modeling procedure, ground test validation, flight test clearance, and test data analysis for the reconfigured F/A-18 aircraft. Multivariable ASE stability margins are calculated from flight data and compared to analytical margins. Because this thrust-vectoring configuration uses exhaust vanes to vector the thrust, the modeling issues are nearly identical for modem multi-axis nozzle configurations. This report correlates analysis results with flight test data and makes observations concerning the application of the linear predictions to thrust-vectoring and high-AOA flight.

Ares I-X Flight Test Vehicle Modal Test

NASA Technical Reports Server (NTRS)

Buehrle, Ralph D.; Templeton, Justin D.; Reaves, Mercedes C.; Horta, Lucas G.; Gaspar, James L.; Bartolotta, Paul A.; Parks, Russel A.; Lazor, Daniel R.

2010-01-01

The first test flight of NASA's Ares I crew launch vehicle, called Ares I-X, was launched on October 28, 2009. Ares I-X used a 4-segment reusable solid rocket booster from the Space Shuttle heritage with mass simulators for the 5th segment, upper stage, crew module and launch abort system. Flight test data will provide important information on ascent loads, vehicle control, separation, and first stage reentry dynamics. As part of hardware verification, a series of modal tests were designed to verify the dynamic finite element model (FEM) used in loads assessments and flight control evaluations. Based on flight control system studies, the critical modes were the first three free-free bending mode pairs. Since a test of the free-free vehicle was not practical within project constraints, modal tests for several configurations during vehicle stacking were defined to calibrate the FEM. Test configurations included two partial stacks and the full Ares I-X flight test vehicle on the Mobile Launcher Platform. This report describes the test requirements, constraints, pre-test analysis, test execution and results for the Ares I-X flight test vehicle modal test on the Mobile Launcher Platform. Initial comparisons between pre-test predictions and test data are also presented.

Postflight hardware evaluation 360T026 (RSRM-26, STS-47)

NASA Technical Reports Server (NTRS)

Nielson, Greg

1993-01-01

The final report for the Clearfield disassembly evaluation and a continuation of the KSC postflight assessment for the 360T026 (STS-47) Redesigned Solid Rocket Motor (RSRM) flight set is provided. All observed hardware conditions were documented on PFOR's and are included in Appendices A, B, and C. Appendices D and E contain the measurements and safety factor data for the nozzle and insulation components. This report, along with the KSC Ten-Day Postflight Hardware Evaluation Report (TWR-64203), represents a summary of the 360T026 hardware evaluation. The as-flown hardware configuration is documented in TWR-60472. Disassembly evaluation photograph numbers are logged in TWA-1987. The 360T026 flight set disassembly evaluations described were performed at the RSRM Refurbishment Facility in Clearfield, Utah. The final factory joint demate occurred on 12 April 1993. Detailed evaluations were performed in accordance with the Clearfield Postflight Engineering Evaluation Plan (PEEP), TWR-50051, Revision A. All observations were compared against limits that are also defined in the PEEP. These limits outline the criteria for categorizing the observations as acceptable, reportable, or critical. Hardware conditions that were unexpected and/or determined to be reportable or critical were evaluated by the applicable CPT and tracked through the PFAR system.

NASA's Space Launch System: A Transformative Capability for Deep Space Missions

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2017-01-01

Already making substantial progress toward its first launches, NASA’s Space Launch System (SLS) exploration-class launch vehicle presents game-changing new opportunities in spaceflight, enabling human exploration of deep space, as well as a variety of missions and mission profiles that are currently impossible. Today, the initial configuration of SLS, able to deliver more than 70 metric tons of payload to low Earth orbit (LEO), is well into final production and testing ahead of its planned first flight, which will send NASA’s new Orion crew vehicle around the moon and will deploy 13 CubeSats, representing multiple disciplines, into deep space. At the same time, production work is already underway toward the more-capable Block 1B configuration, planned to debut on the second flight of SLS, and capable of lofting 105 tons to LEO or of co-manifesting large exploration systems with Orion on launches to the lunar vicinity. Progress being made on the vehicle for that second flight includes initial welding of its core stage and testing of one of its engines, as well as development of new elements such as the powerful Exploration Upper Stage and the Universal Stage Adapter “payload bay.” Ultimately, SLS will evolve to a configuration capable of delivering more than 130 tons to LEO to support humans missions to Mars. In order to enable human deep-space exploration, SLS provides unrivaled mass, volume, and departure energy for payloads, offering numerous benefits for a variety of other missions. For robotic science probes to the outer solar system, for example, SLS can cut transit times to less than half that of currently available vehicles or substantially increased spacecraft mass. In the field of astrophysics, SLS’ high payload volume, in the form of payload fairings with a diameter of up to 10 meters, creates the opportunity for launch of large-aperture telescopes providing an unprecedented look at our universe. This presentation will give an overview of SLS’ capabilities and its current status, and discuss the vehicle’s potential for human exploration of deep space and other game-changing utilization opportunities.

AFTI/F-16

NASA Technical Reports Server (NTRS)

1991-01-01

The AFTI F-16 flying at high angle of attack, shown in the final configuration and paint finish. Dummy Sidewinder air-to-air missles are attached to the wing tips. The white objects visible on the wing racks represent practice bomb dispensers, used in weapon tests. During the 1980s and 1990s, NASA and the U.S. Air Force participated in a joint program to integrate and demonstrate new avionics technologies to improve close air support capabilities in next-generation aircraft. The testbed aircraft, seen here in flight over the desert at NASA's Dryden Flight Research Center, Edwards, California, was called the Advanced Fighter Technology Integration (AFTI) F-16. The tests demonstrated technologies to improve navigation and the pilot's ability to find and destroy enemy ground targets day or night, including adverse weather. The aircraft--an F-16A Fighting Falcon (Serial #75-0750)--underwent numerous modifications. A relatively low-cost testbed, it evaluated the feasability of advanced, intergrated-sensor, avionics, and flight control technologies. During the first phase of the AFTI/F-16 program, which began in 1983, the aircraft demonstrated voice-actuated commands, helmet-mounted sights, flat turns, and selective fuselage pointing using forward-mounted canards and a triplex digital flight control computer system. The second phase of research, which began in the summer of 1991, demonstrated advanced technologies and capabilities to find and destroy ground targets day or night, and in adverse weather while using maneuverability and speed at low altitude. This phase was known as the close air support and battlefield air interdiction (CAS/BAI) phase. Finally, the aircraft was used to assess the Automatic Ground Collision Avoidance System (Auto - GCAS), a joint project with the Swedish Government. For these tests, the pilot flew the aircraft directly toward the ground, simulating a total loss of control. The GCAS was designed to take command in such emergencies and bring the aircraft back to level flight. The AFTI F-16 program ended at Dryden on November 4, 1997 after 15 years and over 700 research flights. The USAF continued to fly the aircraft until retiring it to the Air Force Museum on January 9, 2001.

AFTI/F-16

NASA Technical Reports Server (NTRS)

1992-01-01

The AFTI F-16 in its final configuration, flying in the vicinity of Edwards Air Force Base, California. During this phase, the two forward infrared turrets were added ahead of the cockpit, the chin canards were removed, and the aircraft was repainted in a standard Air Force scheme. A fuel drop tank is visible below the wing. During the 1980s and 1990s, NASA and the U.S. Air Force participated in a joint program to integrate and demonstrate new avionics technologies to improve close air support capabilities in next-generation aircraft. The testbed aircraft, seen here in flight over the desert at NASA's Dryden Flight Research Center, Edwards, California, was called the Advanced Fighter Technology Integration (AFTI) F-16. The tests demonstrated technologies to improve navigation and the pilot's ability to find and destroy enemy ground targets day or night, including adverse weather. The aircraft--an F-16A Fighting Falcon (Serial #75-0750)--underwent numerous modifications. A relatively low-cost testbed, it evaluated the feasability of advanced, intergrated-sensor, avionics, and flight control technologies. During the first phase of the AFTI/F-16 program, which began in 1983, the aircraft demonstrated voice-actuated commands, helmet-mounted sights, flat turns, and selective fuselage pointing using forward-mounted canards and a triplex digital flight control computer system. The second phase of research, which began in the summer of 1991, demonstrated advanced technologies and capabilities to find and destroy ground targets day or night, and in adverse weather while using maneuverability and speed at low altitude. This phase was known as the close air support and battlefield air interdiction (CAS/BAI) phase. Finally, the aircraft was used to assess the Automatic Ground Collision Avoidance System (Auto - GCAS), a joint project with the Swedish Government. For these tests, the pilot flew the aircraft directly toward the ground, simulating a total loss of control. The GCAS was designed to take command in such emergencies and bring the aircraft back to level flight. The AFTI F-16 program ended at Dryden on November 4, 1997 after 15 years and over 700 research flights. The USAF continued to fly the aircraft until retiring it to the Air Force Museum on January 9, 2001.

NASA'S Space Launch System: Progress Toward the Proving Ground

NASA Technical Reports Server (NTRS)

Jackman, Angie; Johnson, Les

2017-01-01

With significant and substantial progress being accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already also underway on preparations for the second flight – using an upgraded version of the vehicle – and beyond. Designed to support human missions into deep space, Space Launch System (SLS), is the most powerful human-rated launch vehicle the United States has ever undertaken, and together with the Orion spacecraft will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. For its second flight, SLS will be upgraded to the more-capable Block 1B configuration. While the Block 1 configuration is capable of delivering more than 70 metric tons to low Earth orbit, the Block 1B vehicle will increase that capability to 105 metric tons. For that flight, the new configuration introduces two major new elements to the vehicle – an Exploration Upper Stage (EUS) that will be used for both ascent and in-space propulsion, and a Universal Stage Adapter (USA) that serves as a “payload bay” for the rocket, allowing the launch of large exploration systems along with the Orion spacecraft. Already, flight hardware is being prepared for the Block 1B vehicle. Beyond the second flight, additional upgrades will be made to the vehicle. The Block 1B vehicle will also be able to launch 8.4-meter-diameter payload fairings, larger than any previously flown, and the Spacecraft Payload Integration and Evolution (SPIE) Element will oversee development and production of those fairings. Ultimately, SLS will be evolved to a Block 2 configuration, which will replace the solid rocket boosters on the Block 1 and 1B vehicles with more powerful boosters, and will be capable of delivering at least 130 metric tons to LEO. The Block 2 vehicle will be capable of launching even larger 10-meter diameter fairings, which will enable human mission of Mars. With these fairings, the Block 1B and 2 configurations of SLS will also be enabling for a wide variety of other payloads. For robotic science probes to the outer solar system, for example, SLS can cut transit times to less than half that of currently available vehicles, producing earlier data return, enhancing iterative exploration, and reducing mission cost and risk. In the field of astrophysics, SLS’ high payload volume, in the form of payload fairings with a diameter of up to 10 meters, creates the opportunity for launch of large-aperture telescopes providing an unprecedented look at our universe, and offers the ability to conduct crewed servicing missions to observatories stationed at locations beyond low Earth orbit. This paper will provide a description of the SLS vehicle, and an overview of the vehicle’s capabilities and utilization potential.

Kuipers configures the GCP in the ATV-3

NASA Image and Video Library

2012-04-06

ISS030-E-210829 (6 April 2012) --- European Space Agency astronaut Andre Kuipers, Expedition 30 flight engineer, configures the Gas Control Panel (GCP) in the Automated Transfer Vehicle (ATV-3) currently docked with the International Space Station.

Kuipers configures the GCP in the ATV-3

NASA Image and Video Library

2012-04-06

ISS030-E-210810 (6 April 2012) --- European Space Agency astronaut Andre Kuipers, Expedition 30 flight engineer, configures the Gas Control Panel (GCP) in the Automated Transfer Vehicle (ATV-3) currently docked with the International Space Station.

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Tomek, W. G.; Hall, R. M.; Wahls, R. A.; Luckring, J. M.; Owens, L. R.

2002-01-01

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

ARC-1984-AC82-0198-26

NASA Image and Video Library

1984-03-12

Sikorsky RSRA - Rotor Systems Research Aircraft (72-002 NASA 741) in helicopter configuration flight. Note: Used in publication in Flight Research at Ames; 57 Years of Development and Validation of Aeronautical Technology NASA SP-1998-3300 fig. 131

Upgrades and Modifications of the NASA Ames HFFAF Ballistic Range

NASA Technical Reports Server (NTRS)

Bogdanoff, David W.; Wilder, Michael C.; Cornelison, Charles J.; Perez, Alfredo J.

2017-01-01

The NASA Ames Hypervelocity Free Flight Aerodynamics Facility ballistic range is described. The various configurations of the shadowgraph stations are presented. This includes the original stations with film and configurations with two different types of digital cameras. Resolution tests for the 3 shadowgraph station configurations are described. The advantages of the digital cameras are discussed, including the immediate availability of the shadowgraphs. The final shadowgraph station configuration is a mix of 26 Nikon cameras and 6 PI-MAX2 cameras. Two types of trigger light sheet stations are described visible and IR. The two gunpowders used for the NASA Ames 6.251.50 light gas guns are presented. These are the Hercules HC-33-FS powder (no longer available) and the St. Marks Powder WC 886 powder. The results from eight proof shots for the two powders are presented. Both muzzle velocities and piston velocities are 5 9 lower for the new St. Marks WC 886 powder than for the old Hercules HC-33-FS powder (no longer available). The experimental and CFD (computational) piston and muzzle velocities are in good agreement. Shadowgraph-reading software that employs template-matching pattern recognition to locate the ballistic-range model is described. Templates are generated from a 3D solid model of the ballistic-range model. The accuracy of the approach is assessed using a set of computer-generated test images.

Propfan test assessment testbed aircraft flutter model test report

NASA Technical Reports Server (NTRS)

Jenness, C. M. J.

1987-01-01

The PropFan Test Assessment (PTA) program includes flight tests of a propfan power plant mounted on the left wind of a modified Gulfstream II testbed aircraft. A static balance boom is mounted on the right wing tip for lateral balance. Flutter analyses indicate that these installations reduce the wing flutter stabilizing speed and that torsional stiffening and the installation of a flutter stabilizing tip boom are required on the left wing for adequate flutter safety margins. Wind tunnel tests of a 1/9th scale high speed flutter model of the testbed aircraft were conducted. The test program included the design, fabrication, and testing of the flutter model and the correlation of the flutter test data with analysis results. Excellent correlations with the test data were achieved in posttest flutter analysis using actual model properties. It was concluded that the flutter analysis method used was capable of accurate flutter predictions for both the (symmetric) twin propfan configuration and the (unsymmetric) single propfan configuration. The flutter analysis also revealed that the differences between the tested model configurations and the current aircraft design caused the (scaled) model flutter speed to be significantly higher than that of the aircraft, at least for the single propfan configuration without a flutter boom. Verification of the aircraft final design should, therefore, be based on flutter predictions made with the test validated analysis methods.

Speed of light demonstration using Doppler beat

NASA Astrophysics Data System (ADS)

Bernal, Luis; Bilbao, Luis

2018-05-01

From an apparatus previously designed for measuring the Doppler shift using a rotating mirror, an improved, versatile version was developed for speed of light demonstrations in a classroom or a teaching laboratory. By adding a second detector and adequate beam-splitter and mirrors, three different configurations are easily assembled. One configuration is used for time-of-flight measurements between a near and a far detector, allowing one to measure the speed of light provided that the path length between detectors is known. Another variation is the interferometric method obtained by superposing the far and near signals in such a way that a minimum of the combined signal is obtained when the time delay makes the signals arrive out of phase by π radians. Finally, the standard Doppler configuration allows the measurement of the frequency beat as a function of the rotation frequency. The main advantages of the apparatus are (a) the experimental setup is simple and completely accessible to undergraduate students, (b) the light is visible, students can see the rays, which, with the use of appropriate screens, can be blocked at any point along their paths, (c) the experiment can take place entirely within the teaching laboratory or demonstration room (using the interferometric method, the shortest distance to the far mirror was as small as 0.5 m), and (d) different configurations can be built, including some economical setups within the budget of teaching laboratories.

Peak-Seeking Optimization of Spanwise Lift Distribution for Wings in Formation Flight

NASA Technical Reports Server (NTRS)

Hanson, Curtis E.; Ryan, Jack

2012-01-01

A method is presented for the in-flight optimization of the lift distribution across the wing for minimum drag of an aircraft in formation flight. The usual elliptical distribution that is optimal for a given wing with a given span is no longer optimal for the trailing wing in a formation due to the asymmetric nature of the encountered flow field. Control surfaces along the trailing edge of the wing can be configured to obtain a non-elliptical profile that is more optimal in terms of minimum combined induced and profile drag. Due to the difficult-to-predict nature of formation flight aerodynamics, a Newton-Raphson peak-seeking controller is used to identify in real time the best aileron and flap deployment scheme for minimum total drag. Simulation results show that the peak-seeking controller correctly identifies an optimal trim configuration that provides additional drag savings above those achieved with conventional anti-symmetric aileron trim.

Mathematical model for lift/cruise fan V/STOL aircraft simulator programming data

NASA Technical Reports Server (NTRS)

Bland, M. P.; Fajfar, B.; Konsewicz, R. K.

1976-01-01

Simulation data are reported for the purpose of programming the flight simulator for advanced aircraft for tests of the lift/cruise fan V/STOL Research Technology Aircraft. These simulation tests are to provide insight into problem areas which are encountered in operational use of the aircraft. A mathematical model is defined in sufficient detail to represent all the necessary pertinent aircraft and system characteristics. The model includes the capability to simulate two basic versions of an aircraft propulsion system: (1) the gas coupled configuration which uses insulated air ducts to transmit power between gas generators and fans in the form of high energy engine exhaust and (2) the mechanically coupled power system which uses shafts, clutches, and gearboxes for power transmittal. Both configurations are modeled such that the simulation can include vertical as well as rolling takeoff and landing, hover, powered lift flight, aerodynamic flight, and the transition between powered lift and aerodynamic flight.

NASA's Space Launch System: Enabling Exploration and Discovery

NASA Technical Reports Server (NTRS)

Schorr, Andrew; Robinson, Kimberly F.; Hitt, David

2017-01-01

As NASA's new Space Launch System (SLS) launch vehicle continues to mature toward its first flight and beyond, so too do the agency's plans for utilization of the rocket. Substantial progress has been made toward the production of the vehicle for the first flight of SLS - an initial "Block 1" configuration capable of delivering more than 70 metric tons (t) to Low Earth Orbit (LEO). That vehicle will be used for an uncrewed integrated test flight, propelling NASA's Orion spacecraft into lunar orbit before it returns safely to Earth. Flight hardware for that launch is being manufactured at facilities around the United States, and, in the case of Orion's service module, beyond. At the same time, production has already begun on the vehicle for the second SLS flight, a more powerful Block 1B configuration capable of delivering more than 105 t to LEO. This configuration will be used for crewed launches of Orion, sending astronauts farther into space than anyone has previously ventured. The 1B configuration will introduce an Exploration Upper Stage, capable of both ascent and in-space propulsion, as well as a Universal Stage Adapter - a payload bay allowing the flight of exploration hardware with Orion - and unprecedentedly large payload fairings that will enable currently impossible spacecraft and mission profiles on uncrewed launches. The Block 1B vehicle will also expand on the initial configuration's ability to deploy CubeSat secondary payloads, creating new opportunities for low-cost access to deep space. Development work is also underway on future upgrades to SLS, which will culminate in about a decade in the Block 2 configuration, capable of delivering 130 t to LEO via the addition of advanced boosters. As the first SLS draws closer to launch, NASA continues to refine plans for the human deep-space exploration it will enable. Planning currently focuses on use of the vehicle to assemble a Deep Space Gateway, which would comprise a habitat in the lunar vicinity allowing astronauts to gain experience living and working in deep space, a testbed for new systems and capabilities needed for exploration beyond, and a departure point for NASA and partners to send missions to other destinations. Assembly of the Gateway would be followed by a Deep Space Transport, which would be a vehicle capable of carrying astronauts farther into our solar system and eventually to Mars. This paper will give an overview of SLS' current status and its capabilities, and discuss current utilization planning.

NASA's Space Launch System: Enabling Exploration and Discovery

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Schorr, Andrew

2017-01-01

As NASA's new Space Launch System (SLS) launch vehicle continues to mature toward its first flight and beyond, so too do the agency's plans for utilization of the rocket. Substantial progress has been made toward the production of the vehicle for the first flight of SLS - an initial "Block 1" configuration capable of delivering more than 70 metric tons (t) to Low Earth Orbit (LEO). That vehicle will be used for an uncrewed integrated test flight, propelling NASA's Orion spacecraft into lunar orbit before it returns safely to Earth. Flight hardware for that launch is being manufactured at facilities around the United States, and, in the case of Orion's service module, beyond. At the same time, production has already begun on the vehicle for the second SLS flight, a more powerful Block 1B configuration capable of delivering more than 105 metric tons to LEO. This configuration will be used for crewed launches of Orion, sending astronauts farther into space than anyone has previously ventured. The 1B configuration will introduce an Exploration Upper Stage, capable of both ascent and in-space propulsion, as well as a Universal Stage Adapter - a payload bay allowing the flight of exploration hardware with Orion - and unprecedentedly large payload fairings that will enable currently impossible spacecraft and mission profiles on uncrewed launches. The Block 1B vehicle will also expand on the initial configuration's ability to deploy CubeSat secondary payloads, creating new opportunities for low-cost access to deep space. Development work is also underway on future upgrades to SLS, which will culminate in about a decade in the Block 2 configuration, capable of delivering 130 metric tons to LEO via the addition of advanced boosters. As the first SLS draws closer to launch, NASA continues to refine plans for the human deep-space exploration it will enable. Planning currently focuses on use of the vehicle to assemble a Deep Space Gateway, which would comprise a habitat in the lunar vicinity allowing astronauts to gain experience living and working in deep space, a testbed for new systems and capabilities needed for exploration beyond, and a departure point for NASA and partners to send missions to other destinations. Assembly of the Gateway would be followed by a Deep Space Transport, which would be a vehicle capable of carrying astronauts farther into our solar system and eventually to Mars. This paper will give an overview of SLS' current status and its capabilities, and discuss current utilization planning.

SEPAC software configuration control plan and procedures, revision 1

NASA Technical Reports Server (NTRS)

1981-01-01

SEPAC Software Configuration Control Plan and Procedures are presented. The objective of the software configuration control is to establish the process for maintaining configuration control of the SEPAC software beginning with the baselining of SEPAC Flight Software Version 1 and encompass the integration and verification tests through Spacelab Level IV Integration. They are designed to provide a simplified but complete configuration control process. The intent is to require a minimum amount of paperwork but provide total traceability of SEPAC software.

Space Operations Center system analysis study extension. Volume 4, book 2: SOC system analysis report

NASA Technical Reports Server (NTRS)

1982-01-01

The Space Operations Center (SOC) orbital space station research missions integration, crew requirements, SOC operations, and configurations are analyzed. Potential research and applications missions and their requirements are described. The capabilities of SOC are compared with user requirements. The SOC/space shuttle and shuttle-derived vehicle flight support operations and SOC orbital operations are described. Module configurations and systems options, SOC/external tank configurations, and configurations for geostationary orbits are described. Crew and systems safety configurations are summarized.

Energy efficient engine fan component detailed design report

NASA Technical Reports Server (NTRS)

Halle, J. E.; Michael, C. J.

1981-01-01

The fan component which was designed for the energy efficient engine is an advanced high performance, single stage system and is based on technology advancements in aerodynamics and structure mechanics. Two fan components were designed, both meeting the integrated core/low spool engine efficiency goal of 84.5%. The primary configuration, envisioned for a future flight propulsion system, features a shroudless, hollow blade and offers a predicted efficiency of 87.3%. A more conventional blade was designed, as a back up, for the integrated core/low spool demonstrator engine. The alternate blade configuration has a predicted efficiency of 86.3% for the future flight propulsion system. Both fan configurations meet goals established for efficiency surge margin, structural integrity and durability.

Control of forward swept wing configurations dominated by flight-dynamic/aeroelastic interactions

NASA Technical Reports Server (NTRS)

Rimer, M.; Chipman, R.; Muniz, B.

1984-01-01

An active control system concept for an aeroelastic wind-tunnel model of a statically unstable FSW configuration with wing-mounted stores is developed to provide acceptable longitudinal flying qualities while maintaining adequate flutter speed margin. On FSW configurations, the inherent aeroelastic wing divergence tendency causes strong flight-dynamic/aeroelastic interactions that in certain cases can produce a dynamic instability known as body-freedom flutter (BFF). The carriage of wing-mounted stores is shown to severely aggravate this problem. The control system developed combines a canard-based SAS with an Active Divergence/Flutter Suppression (ADFS) system which relies on wing-mounted sensors and a trailing-edge device (flaperon). Synergism between these two systems is exploited to obtain the flying qualities and flutter speed objectives.

Flying qualities criteria for GA single pilot IFR operations

NASA Technical Reports Server (NTRS)

Bar-Gill, A.

1982-01-01

The flying qualities criteria in general aviation (GA) to decrease accidents are discussed. The following in-flight research is discussed: (1) identification of key aerodynamic configurations; (2) implementation of an in-flight simulator; (3) mission matrix design; (4) experimental systems; (5) data reduction; (6) optimal flight path reconstruction. Some of the accomplished work is reported: an integrated flight testing and flight path reconstruction methodology was developd, high accuracy in trajectory estimation was achieved with an experimental setup, and a part of the flight test series was flown.

Managing Risk for Thermal Vacuum Testing of the International Space Station Radiators

NASA Technical Reports Server (NTRS)

Carek, Jerry A.; Beach, Duane E.; Remp, Kerry L.

2000-01-01

The International Space Station (ISS) is designed with large deployable radiator panels that are used to reject waste heat from the habitation modules. Qualification testing of the Heat Rejection System (HRS) radiators was performed using qualification hardware only. As a result of those tests, over 30 design changes were made to the actual flight hardware. Consequently, a system level test of the flight hardware was needed to validate its performance in the final configuration. A full thermal vacuum test was performed on the flight hardware in order to demonstrate its ability to deploy on-orbit. Since there is an increased level of risk associated with testing flight hardware, because of cost and schedule limitations, special risk mitigation procedures were developed and implemented for the test program, This paper introduces the Continuous Risk Management process that was utilized for the ISS HRS test program. Testing was performed in the Space Power Facility at the NASA Glenn Research Center, Plum Brook Station located in Sandusky, Ohio. The radiator system was installed in the 100-foot diameter by 122-foot tall vacuum chamber on a special deployment track. Radiator deployments were performed at several thermal conditions similar to those expected on-orbit using both the primary deployment mechanism and the back-up deployment mechanism. The tests were highly successful and were completed without incident.

The development and evaluation of airborne in situ N2O and CH4 sampling using a quantum cascade laser absorption spectrometer (QCLAS)

NASA Astrophysics Data System (ADS)

Pitt, J. R.; Le Breton, M.; Allen, G.; Percival, C. J.; Gallagher, M. W.; Bauguitte, S. J.-B.; O'Shea, S. J.; Muller, J. B. A.; Zahniser, M. S.; Pyle, J.; Palmer, P. I.

2016-01-01

Spectroscopic measurements of atmospheric N2O and CH4 mole fractions were made on board the FAAM (Facility for Airborne Atmospheric Measurements) large atmospheric research aircraft. We present details of the mid-infrared quantum cascade laser absorption spectrometer (QCLAS, Aerodyne Research Inc., USA) employed, including its configuration for airborne sampling, and evaluate its performance over 17 flights conducted during summer 2014. Two different methods of correcting for the influence of water vapour on the spectroscopic retrievals are compared and evaluated. A new in-flight calibration procedure to account for the observed sensitivity of the instrument to ambient pressure changes is described, and its impact on instrument performance is assessed. Test flight data linking this sensitivity to changes in cabin pressure are presented. Total 1σ uncertainties of 2.47 ppb for CH4 and 0.54 ppb for N2O are derived. We report a mean difference in 1 Hz CH4 mole fraction of 2.05 ppb (1σ = 5.85 ppb) between in-flight measurements made using the QCLAS and simultaneous measurements using a previously characterised Fast Greenhouse Gas Analyser (FGGA, Los Gatos Research, USA). Finally, a potential case study for the estimation of a regional N2O flux using a mass balance technique is identified, and the method for calculating such an estimate is outlined.

The development and evaluation of airborne in situ N2O and CH4 sampling using a Quantum Cascade Laser Absorption Spectrometer (QCLAS)

NASA Astrophysics Data System (ADS)

Pitt, J. R.; Le Breton, M.; Allen, G.; Percival, C. J.; Gallagher, M. W.; Bauguitte, S. J.-B.; O'Shea, S. J.; Muller, J. B. A.; Zahniser, M. S.; Pyle, J.; Palmer, P. I.

2015-08-01

Spectroscopic measurements of atmospheric N2O and CH4 mole fractions were made on board the FAAM (Facility for Airborne Atmospheric Measurements) large Atmospheric Research Aircraft. We present details of the mid-IR Aerodyne Research Inc. Quantum Cascade Laser Absorption Spectrometer (QCLAS) employed, including its configuration for airborne sampling, and evaluate its performance over 17 flights conducted during summer 2014. Two different methods of correcting for the influence of water vapour on the spectroscopic retrievals are compared and evaluated. A new in-flight calibration procedure to account for the observed sensitivity of the instrument to ambient pressure changes is described, and its impact on instrument performance is assessed. Test flight data linking this sensitivity to changes in cabin pressure is presented. Total 1σ uncertainties of 1.81 ppb for CH4 and 0.35 ppb for N2O are derived. We report a mean difference in 1 Hz CH4 mole fraction of 2.05 ppb (1σ = 5.85 ppb) between in-flight measurements made using the QCLAS and simultaneous measurements using a previously characterised Los Gatos Research Fast Greenhouse Gas Analyser (FGGA). Finally, a potential case study for the estimation of a regional N2O flux using a mass balance technique is identified, and the method for calculating such an estimate is outlined.

A compact time-of-flight mass spectrometer for ion source characterization

DOE Office of Scientific and Technical Information (OSTI.GOV)

Chen, L., E-mail: l.chen03@gmail.com; Wan, X.; Jin, D. Z.

2015-03-15

A compact time-of-flight mass spectrometer with overall dimension of about 413 × 250 × 414 mm based on orthogonal injection and angle reflection has been developed for ion source characterization. Configuration and principle of the time-of-flight mass spectrometer are introduced in this paper. The mass resolution is optimized to be about 1690 (FWHM), and the ion energy detection range is tested to be between about 3 and 163 eV with the help of electron impact ion source. High mass resolution and compact configuration make this spectrometer useful to provide a valuable diagnostic for ion spectra fundamental research and study themore » mass to charge composition of plasma with wide range of parameters.« less

Supercritical Wing Technology: A Progress Report on Flight Evaluations

NASA Technical Reports Server (NTRS)

1972-01-01

The papers in this compilation were presented at the NASA Symposium on "Supercritical Wing Technology: A Progress Report on Flight Evaluation" held at the NASA Flight Research Center, Edwards, Calif., on February 29, 1972. The purpose of the symposium was to present timely information on flight results obtained with the F-8 and T-2C supercritical wing configurations, discuss comparisons with wind-tunnel predictions, and project [ ] flight programs planned for the F-8 and F-III (TACT) airplanes.

Low-speed wind-tunnel investigation of the flight dynamic characteristics of an advanced turboprop business/commuter aircraft configuration

NASA Technical Reports Server (NTRS)

Coe, Paul L., Jr.; Turner, Steven G.; Owens, D. Bruce

1990-01-01

An investigation was conducted to determine the low-speed flight dynamic behavior of a representative advanced turboprop business/commuter aircraft concept. Free-flight tests were conducted in the NASA Langley Research Center's 30- by 60-Foot Tunnel. In support of the free-flight tests, conventional static, dynamic, and free-to-roll oscillation tests were performed. Tests were intended to explore normal operating and post stall flight conditions, and conditions simulating the loss of power in one engine.

Space tug point design study. Volume 2: Operations, performance and requirements

NASA Technical Reports Server (NTRS)

1973-01-01

A design study to determine the configuration and characteristics of a space tug was conducted. Among the subjects analyzed in the study are: (1) flight and ground operations, (2) vehicle flight performance and performance enhancement techniques, (3) flight requirements, (4) basic design criteria, and (5) functional and procedural interface requirements between the tug and other systems.

Qualification of the flight-critical AFTI/F-16 digital flight control system. [Advanced Fighter Technology Integration

NASA Technical Reports Server (NTRS)

Mackall, D. A.; Ishmael, S. D.; Regenie, V. A.

1983-01-01

Qualification considerations for assuring the safety of a life-critical digital flight control system include four major areas: systems interactions, verification, validation, and configuration control. The AFTI/F-16 design, development, and qualification illustrate these considerations. In this paper, qualification concepts, procedures, and methodologies are discussed and illustrated through specific examples.

MSFC Flight Mission Directive Apollo-Saturn 205 Mission

NASA Technical Reports Server (NTRS)

1966-01-01

The purpose of this directive is to provide, under one cover, coordinated direction for the AS-205 Space Vehicle Flight. Within this document, mission objectives are specified, vehicle configuration is described and referenced, flight trajectories, data acquisition requirements, instrumentation requirements, and detailed documentation requirements necessary to meet launch vehicle mission objectives are defined and/or referenced.

KSC-06pd1159

NASA Image and Video Library

2006-06-16

KENNEDY SPACE CENTER, FLA. - At Astrotech Space Operations in Titusville, Fla., the tilt table lowers the STEREO spacecraft "A." In this position, technicians can perform the final comprehensive performance test of the instruments, verifying the instrument is fully functional before flight. After a rotation, this configuration also allows deployment tests to be done on the solar arrays. STEREO stands for Solar Terrestrial Relations Observatory. The STEREO mission is the first to take measurements of the sun and solar wind in 3-dimension. This new view will improve our understanding of space weather and its impact on the Earth. STEREO is expected to lift off aboard a Boeing Delta II rocket on July 22. Photo credit: NASA/George Shelton

A numerical study of mobility in thin films of fullerene derivatives.

PubMed

Mackenzie, Roderick C I; Frost, Jarvist M; Nelson, Jenny

2010-02-14

The effect of functional group size on the electron mobility in films of fullerene derivatives is investigated numerically. A series of four C(60) derivatives are formed by attaching saturated hydrocarbon chains to the C(60) cage via a methano bridge. For each of the derivatives investigated, molecular dynamics is used to generate a realistic material morphology. Quantum chemical methods are then used to calculate intermolecular charge transfer rates. Finally, Monte Carlo methods are used to simulate time-of-flight experiments and thus calculate the electron mobility. It is found that as the length of the aliphatic side chain increases, the configurational disorder increases and thus the mobility decreases.

Development of flying qualities criteria for single pilot instrument flight operations

NASA Technical Reports Server (NTRS)

Bar-Gill, A.; Nixon, W. B.; Miller, G. E.

1982-01-01

Flying qualities criteria for Single Pilot Instrument Flight Rule (SPIFR) operations were investigated. The ARA aircraft was modified and adapted for SPIFR operations. Aircraft configurations to be flight-tested were chosen and matched on the ARA in-flight simulator, implementing modern control theory algorithms. Mission planning and experimental matrix design were completed. Microprocessor software for the onboard data acquisition system was debugged and flight-tested. Flight-path reconstruction procedure and the associated FORTRAN program were developed. Algorithms associated with the statistical analysis of flight test results and the SPIFR flying qualities criteria deduction are discussed.

Integrated Application of Active Controls (IAAC) technology to an advanced subsonic transport project: Final ACT configuration evaluation

NASA Technical Reports Server (NTRS)

1982-01-01

The Final ACT Configuration Evaluation Task of the Integrated Application of Active Controls (IAAC) technology project within the energy efficient transport program is summarized. The Final ACT Configuration, through application of Active Controls Technology (ACT) in combination with increased wing span, exhibits significant performance improvements over the conventional baseline configuration. At the design range for these configurations, 3590 km, the block fuel used is 10% less for the Final ACT Configuration, with significant reductions in fuel usage at all operational ranges. Results of this improved fuel usage and additional system and airframe costs and the complexity required to achieve it were analyzed to determine its economic effects. For a 926 km mission, the incremental return on investment is nearly 25% at 1980 fuel prices. For longer range missions or increased fuel prices, the return is greater. The technical risks encountered in the Final ACT Configuration design and the research and development effort required to reduce these risks to levels acceptable for commercial airplane design are identified.

Automated Parameter Studies Using a Cartesian Method

NASA Technical Reports Server (NTRS)

Murman, Scott M.; Aftosimis, Michael J.; Nemec, Marian

2004-01-01

Computational Fluid Dynamics (CFD) is now routinely used to analyze isolated points in a design space by performing steady-state computations at fixed flight conditions (Mach number, angle of attack, sideslip), for a fixed geometric configuration of interest. This "point analysis" provides detailed information about the flowfield, which aides an engineer in understanding, or correcting, a design. A point analysis is typically performed using high fidelity methods at a handful of critical design points, e.g. a cruise or landing configuration, or a sample of points along a flight trajectory.

Prediction of forces and moments for flight vehicle control effectors: Workplan

NASA Technical Reports Server (NTRS)

Maughmer, Mark D.

1989-01-01

Two research activities directed at hypersonic vehicle configurations are currently underway. The first involves the validation of a number of classical local surface inclination methods commonly employed in preliminary design studies of hypersonic flight vehicles. Unlike studies aimed at validating such methods for predicting overall vehicle aerodynamics, this effort emphasizes validating the prediction of forces and moments for flight control studies. Specifically, several vehicle configurations for which experimental or flight-test data are available are being examined. By comparing the theoretical predictions with these data, the strengths and weaknesses of the local surface inclination methods can be ascertained and possible improvements suggested. The second research effort, of significance to control during take-off and landing of most proposed hypersonic vehicle configurations, is aimed at determining the change due to ground effect in control effectiveness of highly swept delta planforms. Central to this research is the development of a vortex-lattice computer program which incorporates an unforced trailing vortex sheet and an image ground plane. With this program, the change in pitching moment of the basic vehicle due to ground proximity, and whether or not there is sufficient control power available to trim, can be determined. In addition to the current work, two different research directions are suggested for future study. The first is aimed at developing an interactive computer program to assist the flight controls engineer in determining the forces and moments generated by different types of control effectors that might be used on hypersonic vehicles. The first phase of this work would deal in the subsonic portion of the flight envelope, while later efforts would explore the supersonic/hypersonic flight regimes. The second proposed research direction would explore methods for determining the aerodynamic trim drag of a generic hypersonic flight vehicle and ways in which it can be minimized through vehicle design and trajectory optimization.

StarBooster Demonstrator Cluster Configuration Analysis/Verification Program

NASA Technical Reports Server (NTRS)

DeTurris, Dianne J.

2003-01-01

In order to study the flight dynamics of the cluster configuration of two first stage boosters and upper-stage, flight-testing of subsonic sub-scale models has been undertaken using two glideback boosters launched on a center upper-stage. Three high power rockets clustered together were built and flown to demonstrate vertical launch, separation and horizontal recovery of the boosters. Although the boosters fly to conventional aircraft landing, the centerstage comes down separately under its own parachute. The goal of the project has been to collect data during separation and flight for comparison with a six degree of freedom simulation. The configuration for the delta wing canard boosters comes from a design by Starcraft Boosters, Inc. The subscale rockets were constructed of foam covered in carbon or fiberglass and were launched with commercially available solid rocket motors. The first set of boosters built were 3-ft tall with a 4-ft tall centerstage, and two additional sets of boosters were made that were each over 5-ft tall with a 7.5 ft centerstage. The rocket cluster is launched vertically, then after motor bum out the boosters are separated and flown to a horizontal landing under radio-control. An on-board data acquisition system recorded data during both the launch and glide phases of flight.

Flight Test of an Adaptive Configuration Optimization System for Transport Aircraft

NASA Technical Reports Server (NTRS)

Gilyard, Glenn B.; Georgie, Jennifer; Barnicki, Joseph S.

1999-01-01

A NASA Dryden Flight Research Center program explores the practical application of real-time adaptive configuration optimization for enhanced transport performance on an L-1011 aircraft. This approach is based on calculation of incremental drag from forced-response, symmetric, outboard aileron maneuvers. In real-time operation, the symmetric outboard aileron deflection is directly optimized, and the horizontal stabilator and angle of attack are indirectly optimized. A flight experiment has been conducted from an onboard research engineering test station, and flight research results are presented herein. The optimization system has demonstrated the capability of determining the minimum drag configuration of the aircraft in real time. The drag-minimization algorithm is capable of identifying drag to approximately a one-drag-count level. Optimizing the symmetric outboard aileron position realizes a drag reduction of 2-3 drag counts (approximately 1 percent). Algorithm analysis of maneuvers indicate that two-sided raised-cosine maneuvers improve definition of the symmetric outboard aileron drag effect, thereby improving analysis results and consistency. Ramp maneuvers provide a more even distribution of data collection as a function of excitation deflection than raised-cosine maneuvers provide. A commercial operational system would require airdata calculations and normal output of current inertial navigation systems; engine pressure ratio measurements would be optional.

Ares I-X Launch Vehicle Modal Test Overview

NASA Technical Reports Server (NTRS)

Buehrle, Ralph D.; Bartolotta, Paul A.; Templeton, Justin D.; Reaves, Mercedes C.; Horta, Lucas G.; Gaspar, James L.; Parks, Russell A.; Lazor, Daniel R.

2010-01-01

The first test flight of NASA's Ares I crew launch vehicle, called Ares I-X, is scheduled for launch in 2009. Ares IX will use a 4-segment reusable solid rocket booster from the Space Shuttle heritage with mass simulators for the 5th segment, upper stage, crew module and launch abort system. Flight test data will provide important information on ascent loads, vehicle control, separation, and first stage reentry dynamics. As part of hardware verification, a series of modal tests were designed to verify the dynamic finite element model (FEM) used in loads assessments and flight control evaluations. Based on flight control system studies, the critical modes were the first three free-free bending mode pairs. Since a test of the free-free vehicle is not practical within project constraints, modal tests for several configurations in the nominal integration flow were defined to calibrate the FEM. A traceability study by Aerospace Corporation was used to identify the critical modes for the tested configurations. Test configurations included two partial stacks and the full Ares I-X launch vehicle on the Mobile Launcher Platform. This paper provides an overview for companion papers in the Ares I-X Modal Test Session. The requirements flow down, pre-test analysis, constraints and overall test planning are described.

20 plus Years of Computational Fluid Dynamics for the Space Shuttle

NASA Technical Reports Server (NTRS)

Gomez, Reynaldo J., III

2011-01-01

This slide presentation reviews the use of computational fluid dynamics in performing analysis of the space shuttle with particular reference to the return to flight analysis and other shuttle problems. Slides show a comparison of pressure coefficient with the shuttle ascent configuration between the wind tunnel test and the computed values. the evolution of the grid system for the space shuttle launch vehicle (SSLv) from the early 80's to one in 2004, the grid configuration of the bipod ramp redesign from the original design to the current configuration, charts with the computations showing solid rocket booster surface pressures from wind tunnel data, calculated over two grid systems (i.e., the original 14 grid system, and the enhanced 113 grid system), and the computed flight orbiter wing loads are compared with strain gage data on STS-50 during flight. The loss of STS-107 initiated an unprecedented review of all external environments. The current SSLV grid system of 600+ grids, 1.8 Million surface points and 95+ million volume points is shown. The inflight entry analyses is shown, and the use of Overset CFD as a key part to many external tank redesign and debris assessments is discussed. The work that still remains to be accomplished for future shuttle flights is discussed.

Flight dynamic investigations of flying wing with winglet configured unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Ro, Kapseong

2006-05-01

A swept wing tailless vehicle platform is well known in the radio control (RC) and sailing aircraft community for excellent spiral stability during soaring or thermaling, while exhibiting no Dutch roll behavior at high speed. When an unmanned aerial vehicle (UAV) is subjected to fly a mission in a rugged mountainous terrain where air current or thermal up-drift is frequently present, this is great aerodynamic benefit over the conventional cross-tailed aircraft which requires careful balance between lateral and directional stability. Such dynamic characteristics can be studied through vehicle dynamic modeling and simulation, but it requires configuration aerodynamic data through wind tunnel experiments. Obtaining such data is very costly and time consuming, and it is not feasible especially for low cost and dispensable UAVs. On the other hand, the vehicle autonomy is quite demanding which requires substantial understanding of aircraft dynamic characteristics. In this study, flight dynamics of an UAV platform based on flying wing with a large winglet was investigated through analytical modeling and numerical simulation. Flight dynamic modeling software and experimental formulae were used to obtain essential configuration aerodynamic characteristics, and linear flight dynamic analysis was carried out to understand the effect of wing sweep angle and winglet size on the vehicle dynamic characteristics.

EMC system test performance on Spacelab

NASA Astrophysics Data System (ADS)

Schwan, F.

1982-07-01

Electromagnetic compatibility testing of the Spacelab engineering model is discussed. Documentation, test procedures (including data monitoring and test configuration set up) and performance assessment approach are described. Equipment was assembled into selected representative flight configurations. The physical and functional interfaces between the subsystems were demonstrated within the integration and test sequence which culminated in the flyable configuration Long Module plus one Pallet.

Exploratory Investigation of Aerodynamic Characteristics of Helicopter Tail Boom Cross-Section Models with Passive Venting

NASA Technical Reports Server (NTRS)

Banks, Daniel W.; Kelley, Henry L.

2000-01-01

Two large-scale, two-dimensional helicopter tail boom models were used to determine the effects of passive venting on boom down loads and side forces in hovering crosswind conditions. The models were oval shaped and trapezoidal shaped. Completely porous and solid configurations, partial venting in various symmetric and asymmetric configurations, and strakes were tested. Calculations were made to evaluate the trends of venting and strakes on power required when applied to a UH-60 class helicopter. Compared with the UH-60 baseline, passive venting reduced side force but increased down load at flow conditions representing right sideward flight. Selective asymmetric venting resulted in side force benefits close to the fully porous case. Calculated trends on the effects of venting on power required indicated that the high asymmetric oval configuration was the most effective venting configuration for side force reduction, and the high asymmetric with a single strake was the most effective for overall power reduction. Also, curves of side force versus flow angle were noticeable smoother for the vented configurations compared with the solid baseline configuration; this indicated a potential for smoother flight in low-speed crosswind conditions.

Air Launch: Examining Performance Potential of Various Configurations and Growth Options

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Creech, Dennis M.; Philips, Alan

2013-01-01

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 (?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 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.

High speed research system study. Advanced flight deck configuration effects

NASA Technical Reports Server (NTRS)

Swink, Jay R.; Goins, Richard T.

1992-01-01

In mid-1991 NASA contracted with industry to study the high-speed civil transport (HSCT) flight deck challenges and assess the benefits, prior to initiating their High Speed Research Program (HSRP) Phase 2 efforts, then scheduled for FY-93. The results of this nine-month effort are presented, and a number of the most significant findings for the specified advanced concepts are highlighted: (1) a no nose-droop configuration; (2) a far forward cockpit location; and (3) advanced crew monitoring and control of complex systems. The results indicate that the no nose-droop configuration is critically dependent upon the design and development of a safe, reliable, and certifiable Synthetic Vision System (SVS). The droop-nose configuration would cause significant weight, performance, and cost penalties. The far forward cockpit location, with the conventional side-by-side seating provides little economic advantage; however, a configuration with a tandem seating arrangement provides a substantial increase in either additional payload (i.e., passengers) or potential downsizing of the vehicle with resulting increases in performance efficiencies and associated reductions in emissions. Without a droop nose, forward external visibility is negated and takeoff/landing guidance and control must rely on the use of the SVS. The technologies enabling such capabilities, which de facto provides for Category 3 all-weather operations on every flight independent of weather, represent a dramatic benefits multiplier in a 2005 global ATM network: both in terms of enhanced economic viability and environmental acceptability.

The deep space network, volume 10

NASA Technical Reports Server (NTRS)

1972-01-01

Progress on the Deep Space Network (DSN) supporting research and technology is reported. The objectives, functions and facilities of the DSN are described along with the mission support for the following: interplanetary flight projects, planetary flight projects, and manned space flight projects. Work in advanced engineering and communications systems is reported along with changes in hardware and software configurations in the DSN/MSFN tracking stations.

Northrop Grumman HEC flight coaxial cryocoolers performance

NASA Astrophysics Data System (ADS)

Nguyen, T.; Russo, J.; Basel, G.; Chi, D.; Abelson, L.

2018-05-01

The Northrop Grumman Aerospace Systems (NGAS) has expanded the cryocooler product line to include a single stage High Efficiency Cryocooler (HEC) cooler with a coaxial pulse tube cold head that operates at temperatures down to 45K. The HEC coaxial pulse tube cooler has been adopted by several customers, and has completed acceptance testing to meet program flight requirements. The NGAS TRL 9 HEC is a pulse tube cryocooler with a flexure bearing compressor which has been delivered for a number of flight payloads that are currently operating in space. To date, NGAS has delivered space cryocoolers in several configurations including single stage with a linear cold head and two stage with both linear and coaxial cold heads. The new HEC coaxial cooler uses the same TRL9 HEC compressor with a passive pulse tube cold head, to maintain the flight heritage of the HEC linear cooler. In this paper, we present the flight acceptance test data of the HEC coaxial cryocooler, which includes thermal performance, launch vibration and thermal cycling. The HEC coaxial cooler has demonstrated excellent performance in family with the flight qualified HEC linear cooler. The HEC coaxial cooler provides users with additional flexibility in selecting the cold head configuration to meet their particular applications.

Numerical simulation of divergent rocket-based-combined-cycle performances under the flight condition of Mach 3

NASA Astrophysics Data System (ADS)

Cui, Peng; Xu, WanWu; Li, Qinglian

2018-01-01

Currently, the upper operating limit of the turbine engine is Mach 2+, and the lower limit of the dual-mode scramjet is Mach 4. Therefore no single power systems can operate within the range between Mach 2 + and Mach 4. By using ejector rockets, Rocket-based-combined-cycle can work well in the above scope. As the key component of Rocket-based-combined-cycle, the ejector rocket has significant influence on Rocket-based-combined-cycle performance. Research on the influence of rocket parameters on Rocket-based-combined-cycle in the speed range of Mach 2 + to Mach 4 is scarce. In the present study, influences of Mach number and total pressure of the ejector rocket on Rocket-based-combined-cycle were analyzed numerically. Due to the significant effects of the flight conditions and the Rocket-based-combined-cycle configuration on Rocket-based-combined-cycle performances, flight altitude, flight Mach number, and divergence ratio were also considered. The simulation results indicate that matching lower altitude with higher flight Mach numbers can increase Rocket-based-combined-cycle thrust. For another thing, with an increase of the divergent ratio, the effect of the divergent configuration will strengthen and there is a limit on the divergent ratio. When the divergent ratio is greater than the limit, the effect of divergent configuration will gradually exceed that of combustion on supersonic flows. Further increases in the divergent ratio will decrease Rocket-based-combined-cycle thrust.

Spaceflight Ground Support Equipment Reliability & System Safety Data

NASA Technical Reports Server (NTRS)

Fernandez, Rene; Riddlebaugh, Jeffrey; Brinkman, John; Wilkinson, Myron

2012-01-01

Presented were Reliability Analysis, consisting primarily of Failure Modes and Effects Analysis (FMEA), and System Safety Analysis, consisting of Preliminary Hazards Analysis (PHA), performed to ensure that the CoNNeCT (Communications, Navigation, and Networking re- Configurable Testbed) Flight System was safely and reliably operated during its Assembly, Integration and Test (AI&T) phase. A tailored approach to the NASA Ground Support Equipment (GSE) standard, NASA-STD-5005C, involving the application of the appropriate Requirements, S&MA discipline expertise, and a Configuration Management system (to retain a record of the analysis and documentation) were presented. Presented were System Block Diagrams of selected GSE and the corresponding FMEA, as well as the PHAs. Also discussed are the specific examples of the FMEAs and PHAs being used during the AI&T phase to drive modifications to the GSE (via "redlining" of test procedures, and the placement of warning stickers to protect the flight hardware) before being interfaced to the Flight System. These modifications were necessary because failure modes and hazards were identified during the analysis that had not been properly mitigated. Strict Configuration Management was applied to changes (whether due to upgrades or expired calibrations) in the GSE by revisiting the FMEAs and PHAs to reflect the latest System Block Diagrams and Bill Of Material. The CoNNeCT flight system has been successfully assembled, integrated, tested, and shipped to the launch site without incident. This demonstrates that the steps taken to safeguard the flight system when it was interfaced to the various GSE were successful.

Cooperative control theory and integrated flight and propulsion control

NASA Technical Reports Server (NTRS)

Schmidt, David K.; Schierman, John D.

1994-01-01

This report documents the activities and research results obtained under a grant (NAG3-998) from the NASA Lewis Research Center. The focus of the research was the investigation of dynamic interactions between airframe and engines for advanced ASTOVL aircraft configurations, and the analysis of the implications of these interactions on the stability and performance of the airframe and engine control systems. In addition, the need for integrated flight and propulsion control for such aircraft was addressed. The major contribution of this research was the exposition of the fact that airframe and engine interactions could be present, and their effects could include loss of stability and performance of the control systems. Also, the significance of two directional, as opposed to one-directional, coupling was identified and explained. A multi variable stability and performance analysis methodology was developed, and applied to several candidate aircraft configurations. Also exposed was the fact that with interactions present along with some integrated control approaches, the engine command/limiting logic (which represents an important non-linear component of the engine control system) can impact closed-loop airframe/engine system stability. Finally, a brief investigation of control-law synthesis techniques appropriate for the class of systems was pursued, and it was determined that multi variable techniques, included model-following formulations of LQG and/or H (infinity) methods showed promise. However, for practical reasons, decentralized control architectures are preferred, which is an architecture incompatible with these synthesis methods.

Manual Throttles-Only Control Effectivity for Emergency Flight Control of Transport Aircraft

NASA Technical Reports Server (NTRS)

Stevens, Richard; Burcham, Frank W., Jr.

2009-01-01

If normal aircraft flight controls are lost, emergency flight control may be attempted using only the thrust of engines. Collective thrust is used to control flightpath, and differential thrust is used to control bank angle. One issue is whether a total loss of hydraulics (TLOH) leaves an airplane in a recoverable condition. Recoverability is a function of airspeed, altitude, flight phase, and configuration. If the airplane can be recovered, flight test and simulation results on several transport-class airplanes have shown that throttles-only control (TOC) is usually adequate to maintain up-and-away flight, but executing a safe landing is very difficult. There are favorable aircraft configurations, and also techniques that will improve recoverability and control and increase the chances of a survivable landing. The DHS and NASA have recently conducted a flight and simulator study to determine the effectivity of manual throttles-only control as a way to recover and safely land a range of transport airplanes. This paper discusses TLOH recoverability as a function of conditions, and TOC landability results for a range of transport airplanes, and some key techniques for flying with throttles and making a survivable landing. Airplanes evaluated include the B-747, B-767, B-777, B-757, A320, and B-737 airplanes.

Comparison of Flight Measured, Predicted and Wind Tunnel Measured Winglet Characteristics on a KC-135 Aircraft

NASA Technical Reports Server (NTRS)

Dodson, R. O., Jr.

1982-01-01

One of the objectives of the KC-135 Winglet Flight Research and Demonstration Program was to obtain experimental flight test data to verify the theoretical and wind tunnel winglet aerodynamic performance prediction methods. Good agreement between analytic, wind tunnel and flight test performance was obtained when the known differences between the tests and analyses were accounted for. The flight test measured fuel mileage improvements for a 0.78 Mach number was 3.1 percent at 8 x 10(5) pounds W/delta and 5.5 percent at 1.05 x 10(6) pounds W/delta. Correcting the flight measured data for surface pressure differences between wind tunnel and flight resulted in a fuel mileage improvement of 4.4 percent at 8 x 10(5) pounds W/delta and 7.2 percent at 1.05 x 10(6) pounds W/delta. The performance improvement obtained was within the wind tunnel test data obtained from two different wind tunnel models. The buffet boundary data obtained for the baseline configuration was in good agreement with previous established data. Buffet data for the 15 deg cant/-4 deg incidence configuration showed a slight improvement, while the 15 deg cant/-2 deg incidence and 0 deg cant/-4 deg incidence data showed a slight deterioration.

Multiple spacecraft configuration designs for coordinated flight missions

NASA Astrophysics Data System (ADS)

Fumenti, Federico; Theil, Stephan

2018-06-01

Coordinated flight allows the replacement of a single monolithic spacecraft with multiple smaller ones, based on the principle of distributed systems. According to the mission objectives and to ensure a safe relative motion, constraints on the relative distances need to be satisfied. Initially, differential perturbations are limited by proper orbit design. Then, the induced differential drifts can be properly handled through corrective maneuvers. In this work, several designs are surveyed, defining the initial configuration of a group of spacecraft while counteracting the differential perturbations. For each of the investigated designs, focus is placed upon the number of deployable spacecraft and on the possibility to ensure safe relative motion through station keeping of the initial configuration, with particular attention to the required Δ V budget and the constraints violations.

Analysis of Impulse Load on VEGA SRM Nozzle During Ignition Transient and Effects on TVC Actuators

NASA Astrophysics Data System (ADS)

Fotino, Domenico; Leofanti, Jose Luis; Serraglia, Ferruccio

2012-07-01

During the VEGA development phase and in particular during the Zefiro 23 (second stage motor) on-ground firing tests, values of impulse load on the actuators very close to the requirement were experienced. As a consequence, an activity for the extrapolation of these loads in the flight configuration (longer nozzle and vacuum conditions) was carried out and a mathematical model has been developed with this aim. After providing an overview on the differences between the ground and flight case from the fluid dynamic point of view, the paper describes the results of the mathematical model both in terms of correlation with respect to ground tests and of extrapolation of the loads to the flight configuration. The main effects of this load on the actuators is also addressed.

DC-10 winglet flight evaluation

NASA Technical Reports Server (NTRS)

1983-01-01

The results of a flight evaluation of winglets on a DC-10 Series 10 aircraft are presented. For sensitive areas of comparison, effects of winglets were determined back to back with and without winglets. Basic and reduced span winglet configurations were tested. After initial encounter with low speed buffet, a number of acceptable configurations were developed. For maximum drag reduction at both cruise and low speeds, lower winglets were required, having leading edge devices on upper and lower winglets for the latter regime. The cruise benefits were enhanced by adding outboard aileron droop to the reduced span winglet aircraft. Winglets had no significant impact on stall speeds, high speed buffet boundary, and stability and control characteristics. Flutter test results agreed with predictions and ground vibration data. Flight loads measurement also agreed with predictions.

Mars Science Laboratory Flight Software Boot Robustness Testing Project Report

NASA Technical Reports Server (NTRS)

Roth, Brian

2011-01-01

On the surface of Mars, the Mars Science Laboratory will boot up its flight computers every morning, having charged the batteries through the night. This boot process is complicated, critical, and affected by numerous hardware states that can be difficult to test. The hardware test beds do not facilitate testing a long duration of back-to-back unmanned automated tests, and although the software simulation has provided the necessary functionality and fidelity for this boot testing, there has not been support for the full flexibility necessary for this task. Therefore to perform this testing a framework has been build around the software simulation that supports running automated tests loading a variety of starting configurations for software and hardware states. This implementation has been tested against the nominal cases to validate the methodology, and support for configuring off-nominal cases is ongoing. The implication of this testing is that the introduction of input configurations that have yet proved difficult to test may reveal boot scenarios worth higher fidelity investigation, and in other cases increase confidence in the robustness of the flight software boot process.

System Analyses Driving Improved Aerothermodynamic Lay-Out Of The SpaceLiner Configuration

NASA Astrophysics Data System (ADS)

Sippel, Martin; van Foreest, Arnold; Dietlein, Ingrid; Schwanekamp, Tobias; Kopp, Alexander; Neeb, Dominik

2011-05-01

The revolutionary ultrafast passenger transport SpaceLiner is under investigation at DLR since 2005. The two-stage, fully reusable vehicle is powered by rocket engines. The maximum achieved velocity, depending on the configuration or mission type, is beyond 6.5 km/s putting some challenging aerothermal requirements on the vehicle. At the lower end of the speed-range, the SpaceLiner should have the smallest possible flight velocity for landing with an acceptable angle of attack. The focus of the paper is on all system aspects including the SpaceLiner’s flight performance which have an impact on the aerodynamic configuration. A preliminary sizing of both stages’s passive TPS is done. At the vehicle’s leading edges heat fluxes and hence equilibrium temperatures temporarily reach excessive values requiring advanced active transpiration cooling. An experimental campaign is run at the DLR arc-heated facility to increase the TRL of this promising cooling technology. An aerodynamic shape optimization taking into account trim drag aspects and latest status of the vehicle design and flight profile is described.

Flow visualization of mast-mounted-sight/main rotor aerodynamic interactions

NASA Technical Reports Server (NTRS)

Ghee, Terence A.; Kelley, Henry L.

1993-01-01

Flow visualization tests were conducted on a 27 percent-scale AH-64 attack helicopter model fitted with various mast-mounted-sight configurations in an attempt to identify the cause of adverse vibration encountered during full-scale flight tests of an Apache/Longbow configuration. The tests were conducted at the NASA Langley Research Center in the 14- by 22-Foot Subsonic Tunnel. A symmetric and an asymmetric mast-mounted-sight oriented at several skew angles were tested at forward and rearward flight speeds of 30 and 45 knots. A laser light sheet seeded with vaporized propylene glycol was used to visualize the wake of the sight in planes parallel and perpendicular to the freestream flow. Analysis of the flow visualization data identified the frequency of the wake shed from the sight, the angle-of-attack at the sight, and the location where the sight wake crossed the rotor plane. Differences in wake structure were observed between the various sight configurations and slew angles. Postulations into the cause of the adverse vibration found in flight test are given along with considerations for future tests.

Subsonic tests of an all-flush-pressure-orifice air data system

NASA Technical Reports Server (NTRS)

Larson, T. J.; Siemers, P. M., III

1981-01-01

The use of an all-flush-pressure-orifice array as a subsonic air data system was evaluated in flight and wind tunnel tests. Two orifice configurations were investigated. Both used orifices arranged in a cruciform pattern on the airplane nose. One configuration also used orifices on the sides of the fuselage for a source of static pressure. The all-nose-orifice configuration was similar to the shuttle entry air data system (SEADS). The flight data were obtained with a KC-135A airplane. The wind tunnel data were acquired with a 0.035-scale model of the KC-135A airplane. With proper calibration, several orifices on the vertical centerline of the vehicle's nose were found to be satisfactory for the determination of total pressure and angle of attack. Angle of sideslip could be accurately determined from pressure measurements made on the horizontal centerline of the aircraft. Orifice pairs were also found that provided pressure ratio relationships suitable for the determination of Mach number. The accuracy that can be expected for the air data determined with SEADS during subsonic orbiter flight is indicated.

Recognizing and optimizing flight opportunities with hardware and life sciences limitations.

PubMed

Luttges, M W

1992-01-01

The availability of orbital space flight opportunities to conduct life sciences research has been limited. It is possible to use parabolic flight and sounding rocket programs to conduct some kinds of experiments during short episodes (seconds to minutes) of reduced gravity, but there are constraints and limitations to these programs. Orbital flight opportunities are major undertakings, and the potential science achievable is often a function of the flight hardware available. A variety of generic types of flight hardware have been developed and tested, and show great promise for use during NSTS flights. One such payload configuration is described which has already flown.

Thermal Energy Storage Flight Experiment in Microgravity

NASA Technical Reports Server (NTRS)

Namkoong, David

1992-01-01

The Thermal Energy Storage Flight Experiment was designed to characterize void shape and location in LiF-based phase change materials in different energy storage configurations representative of advanced solar dynamic systems. Experiment goals and payload design are described in outline and graphic form.

B-747 Vortex Alleviation Flight Tests : Ground-Based Sensor Measurements

DOT National Transportation Integrated Search

1982-01-01

In 1979, a series of B-747 flight tests were carried out to study the wake-vortex alleviation produced by deploying spoilers in the landing configuration. The alleviation achieved was examined by encounters of probe aircraft and by velocity profile m...

SPDE: Solar Plasma Diagnostic Experiment

NASA Technical Reports Server (NTRS)

Bruner, Marilyn E.

1995-01-01

The physics of the Solar corona is studied through the use of high resolution soft x-ray spectroscopy and high resolution ultraviolet imagery. The investigation includes the development and application of a flight instrument, first flown in May, 1992 on NASA sounding rocket 36.048. A second flight, NASA founding rocket 36.123, took place on 25 April 1994. Both flights were successful in recording new observations relevant to the investigation. The effort in this contract covers completion of the modifications to the existing rocket payload, its reflight, and the preliminary day reduction and analysis. Experience gained from flight 36.048 led us to plan several payload design modifications. These were made to improve the sensitivity balance between the UV and EUV spectrographs, to improve the scattered light rejection in the spectrographs, to protect the visible light rejection filter for the Normal Incidence X-ray Imager instrument (NIXI), and to prepare one new multilayer mirror coating to the NIXI. We also investigated the addition of a brassboard CCD camera to the payload to test it as a possible replacement for the Eastman type 101-07 film used by the SPDE instruments. This camera was included in the experimeter's data package for the Project Initiation Conference for the flight of NASA Mission 36.123, held in January, 1994, but for programmatic reasons was deleted from the final payload configuration. The payload was shipped to the White Sands Missile Range on schedule in early April. The launch and successful recovery took place on 25 April, in coordination with the Yohkoh satellite and a supporting ground-based observing campaign.

An Entry Flight Controls Analysis for a Reusable Launch Vehicle

NASA Technical Reports Server (NTRS)

Calhoun, Philip

2000-01-01

The NASA Langley Research Center has been performing studies to address the feasibility of various single-stage to orbit concepts for use by NASA and the commercial launch industry to provide a lower cost access to space. Some work on the conceptual design of a typical lifting body concept vehicle, designated VentureStar(sup TM) has been conducted in cooperation with the Lockheed Martin Skunk Works. This paper will address the results of a preliminary flight controls assessment of this vehicle concept during the atmospheric entry phase of flight. The work includes control analysis from hypersonic flight at the atmospheric entry through supersonic speeds to final approach and landing at subsonic conditions. The requirements of the flight control effectors are determined over the full range of entry vehicle Mach number conditions. The analysis was performed for a typical maximum crossrange entry trajectory utilizing angle of attack to limit entry heating and providing for energy management, and bank angle to modulation of the lift vector to provide downrange and crossrange capability to fly the vehicle to a specified landing site. Sensitivity of the vehicle open and closed loop characteristics to CG location, control surface mixing strategy and wind gusts are included in the results. An alternative control surface mixing strategy utilizing a reverse aileron technique demonstrated a significant reduction in RCS torque and fuel required to perform bank maneuvers during entry. The results of the control analysis revealed challenges for an early vehicle configuration in the areas of hypersonic pitch trim and subsonic longitudinal controllability.

Airframe Noise from a Hybrid Wing Body Aircraft Configuration

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Spalt, Taylor B.; Brooks, Thomas F.; Plassman, Gerald E.

2016-01-01

A high fidelity aeroacoustic test was conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel to establish a detailed database of component noise for a 5.8% scale HWB aircraft configuration. The model has a modular design, which includes a drooped and a stowed wing leading edge, deflectable elevons, twin verticals, and a landing gear system with geometrically scaled wheel-wells. The model is mounted inverted in the test section and noise measurements are acquired at different streamwise stations from an overhead microphone phased array and from overhead and sideline microphones. Noise source distribution maps and component noise spectra are presented for airframe configurations representing two different approach flight conditions. Array measurements performed along the aircraft flyover line show the main landing gear to be the dominant contributor to the total airframe noise, followed by the nose gear, the inboard side-edges of the LE droop, the wing tip/LE droop outboard side-edges, and the side-edges of deployed elevons. Velocity dependence and flyover directivity are presented for the main noise components. Decorrelation effects from turbulence scattering on spectral levels measured with the microphone phased array are discussed. Finally, noise directivity maps obtained from the overhead and sideline microphone measurements for the landing gear system are provided for a broad range of observer locations.

Design optimization of high-speed proprotor aircraft

NASA Technical Reports Server (NTRS)

Schleicher, David R.; Phillips, James D.; Carbajal, Kevin B.

1993-01-01

NASA's high-speed rotorcraft (HSRC) studies have the objective of investigating technology for vehicles that have both low downwash velocities and forward flight speed capability of up to 450 knots. This paper investigates a tilt rotor, a tilt wing, and a folding tilt rotor designed for a civil transport mission. Baseline aircraft models using current technology are developed for each configuration using a vertical/short takeoff and landing (V/STOL) aircraft design synthesis computer program to generate converged vehicle designs. Sensitivity studies and numerical optimization are used to illustrate each configuration's key design tradeoffs and constraints. Minimization of the gross takeoff weight is used as the optimization objective function. Several advanced technologies are chosen, and their relative impact on future configurational development is discussed. Finally, the impact of maximum cruise speed on vehicle figures of merit (gross weight, productivity, and direct operating cost) is analyzed. The three most important conclusions from the study are payload ratios for these aircraft will be commensurate with current fixed-wing commuter aircraft; future tilt rotors and tilt wings will be significantly lighter, more productive, and cheaper than competing folding tilt rotors; and the most promising technologies are an advanced-technology proprotor for both tilt rotor and tilt wing and advanced structural materials for the folding tilt rotor.

An atomic gravitational wave interferometric sensor in low earth orbit (AGIS-LEO)

NASA Astrophysics Data System (ADS)

Hogan, Jason M.; Johnson, David M. S.; Dickerson, Susannah; Kovachy, Tim; Sugarbaker, Alex; Chiow, Sheng-Wey; Graham, Peter W.; Kasevich, Mark A.; Saif, Babak; Rajendran, Surjeet; Bouyer, Philippe; Seery, Bernard D.; Feinberg, Lee; Keski-Kuha, Ritva

2011-07-01

We propose an atom interferometer gravitational wave detector in low Earth orbit (AGIS-LEO). Gravitational waves can be observed by comparing a pair of atom interferometers separated by a 30 km baseline. In the proposed configuration, one or three of these interferometer pairs are simultaneously operated through the use of two or three satellites in formation flight. The three satellite configuration allows for the increased suppression of multiple noise sources and for the detection of stochastic gravitational wave signals. The mission will offer a strain sensitivity of {<10^{-18}/sqrt{Hz}} in the 50mHz-10Hz frequency range, providing access to a rich scientific region with substantial discovery potential. This band is not currently addressed with the LIGO, VIRGO, or LISA instruments. We analyze systematic backgrounds that are relevant to the mission and discuss how they can be mitigated at the required levels. Some of these effects do not appear to have been considered previously in the context of atom interferometry, and we therefore expect that our analysis will be broadly relevant to atom interferometric precision measurements. Finally, we present a brief conceptual overview of shorter-baseline ({lesssim100 m}) atom interferometer configurations that could be deployed as proof-of-principle instruments on the International Space Station (AGIS-ISS) or an independent satellite.

The Future of New Discoveries on the International Space Station

NASA Technical Reports Server (NTRS)

Schlagheck, Ronald; Trach, Brian

2000-01-01

The Materials Science program is one of the five Microgravity research disciplines in NASA's Human Exploration and Development of Space (HEDS). This research uses the low gravity environment to obtain the fundamental understanding of various phenomena effects and it's relationship to structure, processing, and properties of materials. The International Space Station (ISS) will complete the first major assembly phase within the next year thus providing the opportunity for on-orbit research and scientific utilization in early 2001. Research will become routine as the final Space Station configuration is completed. Accommodations will support a variety of Materials Science payload hardware both in the US and international partner modules. This paper addresses the current scope of the flight investigator program that will utilize the various capabilities on ISS. The type of research and classification of materials that are addressed using multiple types of flight apparatus will be explained. The various flight and ground facilities that are used to support the NASA program are described. The early utilization schedule for the materials science payloads with associated hardware will be covered. The Materials Science Research Facility and related international experiment modules serves as the foundation for this capability. The potential applications and technologies obtained from the Materials Science program are described.

James Webb Space Telescope (JWST) Integrated Science Instruments Module (ISIM) Cryo-Vacuum (CV) Test Campaign Summary

NASA Technical Reports Server (NTRS)

Yew, Calinda; Lui, Yan; Whitehouse, Paul; Banks, Kimberly

2016-01-01

JWST Integrated Science Instruments Module (ISIM) completed its system-level space simulation testing program at the NASA Goddard Space Flight Center (GSFC). In March 2016, ISIM was successfully delivered to the next level of integration with the Optical Telescope Element (OTE), to form OTIS (OTE + ISIM), after concluding a series of three cryo-vacuum (CV) tests. During these tests, the complexity of the mission has generated challenging requirements that demand highly reliable system performance and capabilities from the Space Environment Simulator (SES) vacuum chamber. The first test served as a risk reduction test; the second test provided the initial verification of the fully-integrated flight instruments; and the third test verified the system in its final flight configuration following mechanical environmental tests (vibration and acoustics). From one test to the next, shortcomings of the facility were uncovered and associated improvements in operational capabilities and reliability of the facility were required to enable the project to verify system-level requirements. This paper: (1) provides an overview of the integrated mechanical and thermal facility systems required to achieve the objectives of JWST ISIM testing, (2) compares the overall facility performance and instrumentation results from the three ISIM CV tests, and (3) summarizes lessons learned from the ISIM testing campaign.

The multipurpose time-of-flight neutron reflectometer “Platypus” at Australia's OPAL reactor

NASA Astrophysics Data System (ADS)

James, M.; Nelson, A.; Holt, S. A.; Saerbeck, T.; Hamilton, W. A.; Klose, F.

2011-03-01

In this manuscript we describe the major components of the Platypus time-of-flight neutron reflectometer at the 20 MW OPAL reactor in Sydney, Australia. Platypus is a multipurpose spectrometer for the characterisation of solid thin films, materials adsorbed at the solid-liquid interface and free-liquid surfaces. It also has the capacity to study magnetic thin films using spin-polarised neutrons. Platypus utilises a white neutron beam ( λ=2-20 Å) that is pulsed using boron-coated disc chopper pairs; thus providing the capacity to tailor the wavelength resolution of the pulses to suit the system under investigation. Supermirror optical components are used to focus, deflect or spin-polarise the broad bandwidth neutron beams, and typical incident spectra are presented for each configuration. A series of neutron reflectivity datasets are presented, indicating the quality and flexibility of this spectrometer. Minimum reflectivity values of <10 -7 are observed; while maximum thickness values of 325 nm have been measured for single-component films and 483 nm for a multilayer system. Off-specular measurements have also been made to investigate in-plane features as opposed to those normal to the sample surface. Finally, the first published studies conducted using the Platypus time-of-flight neutron reflectometer are presented.

SHEFEX II - Aerodynamic Re-Entry Controlled Sharp Edge Flight Experiment

NASA Astrophysics Data System (ADS)

Longo, J. M. A.; Turner, J.; Weihs, H.

2009-01-01

In this paper the basic goals and architecture of the SHEFEX II mission is presented. Also launched by a two staged sounding rocket system SHEFEX II is a consequent next step in technology test and demonstration. Considering all experience and collected flight data obtained during the SHEFEX I Mission, the test vehicle has been re-designed and extended by an active control system, which allows active aerodynamic control during the re-entry phase. Thus, ceramic based aerodynamic control elements like rudders, ailerons and flaps, mechanical actuators and an automatic electronic control unit has been implemented. Special focus is taken on improved GNC Elements. In addition, some other experiments including an actively cooled thermal protection element, advanced sensor equipment, high temperature antenna inserts etc. are part of the SHEFEX II experimental payload. A final 2 stage configuration has been selected considering Brazilian solid rocket boosters derived from the S 40 family. During the experiment phase a maximum entry velocity of Mach around 10 is expected for 50 seconds. Considering these flight conditions, the heat loads are not representative for a RLV re-entry, however, it allows to investigate the principal behaviour of such a facetted ceramic TPS, a sharp leading edge at the canards and fins and all associated gas flow effects and their structural response.

Reduced order model of a blended wing body aircraft configuration

NASA Astrophysics Data System (ADS)

Stroscher, F.; Sika, Z.; Petersson, O.

2013-12-01

This paper describes the full development process of a numerical simulation model for the ACFA2020 (Active Control for Flexible 2020 Aircraft) blended wing body (BWB) configuration. Its requirements are the prediction of aeroelastic and flight dynamic response in time domain, with relatively small model order. Further, the model had to be parameterized with regard to multiple fuel filling conditions, as well as flight conditions. High efforts have been conducted in high-order aerodynamic analysis, for subsonic and transonic regime, by several project partners. The integration of the unsteady aerodynamic databases was one of the key issues in aeroelastic modeling.

In-flight neutron spectra as an ICF diagnostic for implosion asymmetries

NASA Astrophysics Data System (ADS)

Cerjan, C.; Sayre, D. B.; Sepke, S. M.

2018-02-01

The yield and spectral shape of the neutrons produced during in-flight reactions provide stringent constraints upon the symmetry of the fully compressed fuel conditions in Inertial Confinement Fusion implosions. Neutron production from a specific deuterium gas-filled implosion is simulated in detail and compared with the experimental neutron spectra along two lines-of-sight. An approximate reactivity formulation is applied to obtain further insight into the underlying fuel configuration. This analysis suggests that the differences observed in the observed spectra correspond to angularly dependent triton velocity distributions created by an asymmetric plasma configuration.

Simulator study of the stall departure characteristics of a light general aviation airplane with and without a wing-leading-edge modification

NASA Technical Reports Server (NTRS)

Riley, D. R.

1985-01-01

A six-degree-of-freedom nonlinear simulation was developed for a two-place, single-engine, low-wing general aviation airplane for the stall and initial departure regions of flight. Two configurations, one with and one without an outboard wing-leading-edge modification, were modeled. The math models developed are presented simulation predictions and flight-test data for validation purposes and simulation results for the two configurations for various maneuvers and power settings are compared to show the beneficial influence of adding the wing-leading-edge modification.

Early Program Development

NASA Image and Video Library

1961-05-01

This artist's concept illustrates the Module Nova concept - Solid C-3 Basis. From 1960 to 1962, the Marshall Space Flight Center considered the Nova launch vehicle as a means to achieve a marned lunar landing with a direct flight to the Moon. Various configurations of the vehicle were examined. The latest configuration was a five-stage vehicle using eight F-1 engines in the first stage. Although the program was canceled after NASA planners selected the lunar/orbital rendezvous mode, the proposed F-1 engine would eventually be used in the Apollo Program to propel the first stage of the Saturn V launch vehicle.

Early Program Development

NASA Image and Video Library

1961-11-01

This artist's concept illustrates the Module Nova concept - Solid C-3 Basis. From 1960 to 1962, the Marshall Space Flight Center considered the Nova launch vehicle as a means to achieve a marned lunar landing with a direct flight to the Moon. Various configurations of the vehicle were examined. The latest configuration was a five-stage vehicle using eight F-1 engines in the first stage. Although the program was canceled after NASA planners selected the lunar/orbital rendezvous mode, the proposed F-1 engine would eventually be used in the Apollo Program to propel the first stage of the Saturn V launch vehicle.

Status and trends in active control technology

NASA Technical Reports Server (NTRS)

Rediess, H. A.; Szalai, K. J.

1975-01-01

The emergence of highly reliable fly-by-wire flight control systems makes it possible to consider a strong reliance on automatic control systems in the design optimization of future aircraft. This design philosophy has been referred to as the control configured vehicle approach or the application of active control technology. Several studies and flight tests sponsored by the Air Force and NASA have demonstrated the potential benefits of control configured vehicles and active control technology. The present status and trends of active control technology are reviewed and the impact it will have on aircraft designs, design techniques, and the designer is predicted.

Investigation of Aerodynamics Scale Effects for a Generic Fighter Configuration in the National Transonic Facility (Invited)

NASA Technical Reports Server (NTRS)

Tomek, W. G.; Wahls, R. A.; Owens, L. R.; Burner, A. B.; Graves, S. S.; Luckring, J. M.

2003-01-01

Two wind tunnel tests of a generic fighter configuration have been completed in the National Transonic Facility. The primary purpose of the tests was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The tests included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: 1) Fuselage / Wing, 2) Fuselage / Wing / Centerline Vertical Tail / Horizontal Tail, and 3) Fuselage / Wing / Trailing-Edge Extension / Twin Vertical Tails. Reynolds number effects on the lateral-directional aerodynamic characteristics are presented herein, along with longitudinal data demonstrating the effects of fixing the boundary layer transition location for low Reynolds number conditions. In addition, an improved model videogrammetry system and results are discussed.

Flight investigation of the effect of tail configuration on stall, spin, and recovery characteristics of a low-wing general aviation research airplane

NASA Technical Reports Server (NTRS)

Stough, H. Paul, III; Patton, James M., Jr.; Sliwa, Steven M.

1987-01-01

Flight tests were performed to investigate the stall, spin, and recovery characteristics of a low-wing, single-engine, light airplane with four interchangeable tail configurations. The four tail configurations were evaluated for effects of varying mass distribution, center-of-gravity position, and control inputs. The airplane tended to roll-off at the stall. Variations in tail configuration produced spins ranging from 40 deg to 60 deg angle of attack and turn rates of about 145 to 208 deg/sec. Some unrecoverable flat spins were encountered which required use of the airplane spin chute for recovery. For recoverable spins, antispin rudder followed by forward wheel with ailerons centered provided the quickest spin recovery. The moderate spin modes agreed very well with those predicted from spin-tunnel model tests, however, the flat spin was at a lower angle of attack and a slower rotation rate than indicated by the model tests.

Coupled nonlinear aeroelasticity and flight dynamics of fully flexible aircraft

NASA Astrophysics Data System (ADS)

Su, Weihua

This dissertation introduces an approach to effectively model and analyze the coupled nonlinear aeroelasticity and flight dynamics of highly flexible aircraft. A reduced-order, nonlinear, strain-based finite element framework is used, which is capable of assessing the fundamental impact of structural nonlinear effects in preliminary vehicle design and control synthesis. The cross-sectional stiffness and inertia properties of the wings are calculated along the wing span, and then incorporated into the one-dimensional nonlinear beam formulation. Finite-state unsteady subsonic aerodynamics is used to compute airloads along lifting surfaces. Flight dynamic equations are then introduced to complete the aeroelastic/flight dynamic system equations of motion. Instead of merely considering the flexibility of the wings, the current work allows all members of the vehicle to be flexible. Due to their characteristics of being slender structures, the wings, tail, and fuselage of highly flexible aircraft can be modeled as beams undergoing three dimensional displacements and rotations. New kinematic relationships are developed to handle the split beam systems, such that fully flexible vehicles can be effectively modeled within the existing framework. Different aircraft configurations are modeled and studied, including Single-Wing, Joined-Wing, Blended-Wing-Body, and Flying-Wing configurations. The Lagrange Multiplier Method is applied to model the nodal displacement constraints at the joint locations. Based on the proposed models, roll response and stability studies are conducted on fully flexible and rigidized models. The impacts of the flexibility of different vehicle members on flutter with rigid body motion constraints, flutter in free flight condition, and roll maneuver performance are presented. Also, the static stability of the compressive member of the Joined-Wing configuration is studied. A spatially-distributed discrete gust model is incorporated into the time simulation of the framework. Gust responses of the Flying-Wing configuration subject to stall effects are investigated. A bilinear torsional stiffness model is introduced to study the skin wrinkling due to large bending curvature of the Flying-Wing. The numerical studies illustrate the improvements of the existing reduced-order formulation with new capabilities of both structural modeling and coupled aeroelastic and flight dynamic analysis of fully flexible aircraft.

ADEPT Sounding Rocket One (SR-1)Flight Experiment Overview

NASA Technical Reports Server (NTRS)

Wercinski, Paul; Smith, B.; Yount, B.; Cassell, A.; Kruger, C.; Brivkalns, C.; Makino, A.; Duttta, S.; Ghassemieh, S.; Wu, S.;

Space Shuttle Main Engine: Thirty Years of Innovation

NASA Technical Reports Server (NTRS)

Jue, F. H.; Hopson, George (Technical Monitor)

2002-01-01

The Space Shuttle Main Engine (SSME) is the first reusable, liquid booster engine designed for human space flight. This paper chronicles the 30-year history and achievements of the SSME from authority to proceed up to the latest flight configuration - the Block 2 SSME.

Test Results of the Modified Space Shuttle Main Engine at the Marshall Space Flight Center Technology Test Bed Facility

NASA Technical Reports Server (NTRS)

Cook, J.; Dumbacher, D.; Ise, M.; Singer, C.

1990-01-01

A modified space shuttle main engine (SSME), which primarily includes an enlarged throat main combustion chamber with the acoustic cavities removed and a main injector with the stability control baffles removed, was tested. This one-of-a-kind engine's design changes are being evaluated for potential incorporation in the shuttle flight program in the mid-1990's. Engine testing was initiated on September 15, 1988 and has accumulated 1,915 seconds and 19 starts. Testing is being conducted to characterize the engine system performance, combustion stability with the baffle-less injector, and both low pressure oxidizer turbopump (LPOTP) and high pressure oxidizer turbopump (HPOTP) for suction performance. These test results are summarized and compared with the SSME flight configuration data base. Testing of this new generation SSME is the first product from the technology test bed (TTB). Figure test plans for the TTB include the highly instrumented flight configuration SSME and advanced liquid propulsion technology items.

Validation results of satellite mock-up capturing experiment using nets

NASA Astrophysics Data System (ADS)

Medina, Alberto; Cercós, Lorenzo; Stefanescu, Raluca M.; Benvenuto, Riccardo; Pesce, Vincenzo; Marcon, Marco; Lavagna, Michèle; González, Iván; Rodríguez López, Nuria; Wormnes, Kjetil

2017-05-01

The PATENDER activity (Net parametric characterization and parabolic flight), funded by the European Space Agency (ESA) via its Clean Space initiative, was aiming to validate a simulation tool for designing nets for capturing space debris. This validation has been performed through a set of different experiments under microgravity conditions where a net was launched capturing and wrapping a satellite mock-up. This paper presents the architecture of the thrown-net dynamics simulator together with the set-up of the deployment experiment and its trajectory reconstruction results on a parabolic flight (Novespace A-310, June 2015). The simulator has been implemented within the Blender framework in order to provide a highly configurable tool, able to reproduce different scenarios for Active Debris Removal missions. The experiment has been performed over thirty parabolas offering around 22 s of zero-g conditions. Flexible meshed fabric structure (the net) ejected from a container and propelled by corner masses (the bullets) arranged around its circumference have been launched at different initial velocities and launching angles using a pneumatic-based dedicated mechanism (representing the chaser satellite) against a target mock-up (the target satellite). High-speed motion cameras were recording the experiment allowing 3D reconstruction of the net motion. The net knots have been coloured to allow the images post-process using colour segmentation, stereo matching and iterative closest point (ICP) for knots tracking. The final objective of the activity was the validation of the net deployment and wrapping simulator using images recorded during the parabolic flight. The high-resolution images acquired have been post-processed to determine accurately the initial conditions and generate the reference data (position and velocity of all knots of the net along its deployment and wrapping of the target mock-up) for the simulator validation. The simulator has been properly configured according to the parabolic flight scenario, and executed in order to generate the validation data. Both datasets have been compared according to different metrics in order to perform the validation of the PATENDER simulator.

Atmosphere, Magnetosphere and Plasmas in Space (AMPS). Spacelab payload definition study. Volume 2: Mission support requirements document. Addendum: Flight 2

NASA Technical Reports Server (NTRS)

1976-01-01

The AMPS Flight 2 payload, its operation, and the support required from the Space Transportation System (STS) are described. The definition of the payload includes the flight objectives and requirements, the experiment operations, and the payload configuration. The support required from the STS includes the accommodation of the payload by the orbiter/Spacelab, use of the flight operations network and ground facilities, and the use of the launch site facilities.

ER-2 High Altitude Solar Cell Calibration Flights

NASA Technical Reports Server (NTRS)

Piszczor, Michael F.

2017-01-01

ER-2 Capability: Flights can be conducted once every one to two days during a campaign. Flight season runs from April through September. Twelve 2x2cm cells can be flown per flight, or any other configuration that fits inside of the 14.2x14.2cm illuminated area. This capacity could be doubled if the second ER-2 pod is used. Data supplied includes Isc, Voc, IV curve, and cell temperature. Other optical or atmospheric sensors can be flown as able.

Pilot-Configurable Information on a Display Unit

NASA Technical Reports Server (NTRS)

Bell, Charles Frederick (Inventor); Ametsitsi, Julian (Inventor); Che, Tan Nhat (Inventor); Shafaat, Syed Tahir (Inventor)

2017-01-01

A small thin display unit that can be installed in the flight deck for displaying only flight crew-selected tactical information needed for the task at hand. The flight crew can select the tactical information to be displayed by means of any conventional user interface. Whenever the flight crew selects tactical information for processes the request, including periodically retrieving measured current values or computing current values for the requested tactical parameters and returning those current tactical parameter values to the display unit for display.

Iced Aircraft Flight Data for Flight Simulator Validation

NASA Technical Reports Server (NTRS)

Ratvasky, Thomas P.; Blankenship, Kurt; Rieke, William; Brinker, David J.

2003-01-01

NASA is developing and validating technology to incorporate aircraft icing effects into a flight training device concept demonstrator. Flight simulation models of a DHC-6 Twin Otter were developed from wind tunnel data using a subscale, complete aircraft model with and without simulated ice, and from previously acquired flight data. The validation of the simulation models required additional aircraft response time histories of the airplane configured with simulated ice similar to the subscale model testing. Therefore, a flight test was conducted using the NASA Twin Otter Icing Research Aircraft. Over 500 maneuvers of various types were conducted in this flight test. The validation data consisted of aircraft state parameters, pilot inputs, propulsion, weight, center of gravity, and moments of inertia with the airplane configured with different amounts of simulated ice. Emphasis was made to acquire data at wing stall and tailplane stall since these events are of primary interest to model accurately in the flight training device. Analyses of several datasets are described regarding wing and tailplane stall. Key findings from these analyses are that the simulated wing ice shapes significantly reduced the C , max, while the simulated tail ice caused elevator control force anomalies and tailplane stall when flaps were deflected 30 deg or greater. This effectively reduced the safe operating margins between iced wing and iced tail stall as flap deflection and thrust were increased. This flight test demonstrated that the critical aspects to be modeled in the icing effects flight training device include: iced wing and tail stall speeds, flap and thrust effects, control forces, and control effectiveness.

The SOFIA Mission Control System Software

NASA Astrophysics Data System (ADS)

Heiligman, G. M.; Brock, D. R.; Culp, S. D.; Decker, P. H.; Estrada, J. C.; Graybeal, J. B.; Nichols, D. M.; Paluzzi, P. R.; Sharer, P. J.; Pampell, R. J.; Papke, B. L.; Salovich, R. D.; Schlappe, S. B.; Spriestersbach, K. K.; Webb, G. L.

1999-05-01

The Stratospheric Observatory for Infrared Astronomy (SOFIA) will be delivered with a computerized mission control system (MCS). The MCS communicates with the aircraft's flight management system and coordinates the operations of the telescope assembly, mission-specific subsystems, and the science instruments. The software for the MCS must be reliable and flexible. It must be easily usable by many teams of observers with widely differing needs, and it must support non-intrusive access for education and public outreach. The technology must be appropriate for SOFIA's 20-year lifetime. The MCS software development process is an object-oriented, use case driven approach. The process is iterative: delivery will be phased over four "builds"; each build will be the result of many iterations; and each iteration will include analysis, design, implementation, and test activities. The team is geographically distributed, coordinating its work via Web pages, teleconferences, T.120 remote collaboration, and CVS (for Internet-enabled configuration management). The MCS software architectural design is derived in part from other observatories' experience. Some important features of the MCS are: * distributed computing over several UNIX and VxWorks computers * fast throughput of time-critical data * use of third-party components, such as the Adaptive Communications Environment (ACE) and the Common Object Request Broker Architecture (CORBA) * extensive configurability via stored, editable configuration files * use of several computer languages so developers have "the right tool for the job". C++, Java, scripting languages, Interactive Data Language (from Research Systems, Int'l.), XML, and HTML will all be used in the final deliverables. This paper reports on work in progress, with the final product scheduled for delivery in 2001. This work was performed for Universities Space Research Association for NASA under contract NAS2-97001.

Method for Controlling Space Transportation System Life Cycle Costs

NASA Technical Reports Server (NTRS)

McCleskey, Carey M.; Bartine, David E.

2006-01-01

A structured, disciplined methodology is required to control major cost-influencing metrics of space transportation systems during design and continuing through the test and operations phases. This paper proposes controlling key space system design metrics that specifically influence life cycle costs. These are inclusive of flight and ground operations, test, and manufacturing and infrastructure. The proposed technique builds on today's configuration and mass properties control techniques and takes on all the characteristics of a classical control system. While the paper does not lay out a complete math model, key elements of the proposed methodology are explored and explained with both historical and contemporary examples. Finally, the paper encourages modular design approaches and technology investments compatible with the proposed method.

Recent NASA/GSFC cryogenic measurements of the total hemispheric emissivity of black surface preparations

NASA Astrophysics Data System (ADS)

Tuttle, J.; Canavan, E.

2015-12-01

High-emissivity (black) surfaces are commonly used on deep-space radiators and thermal radiation absorbers in test chambers. Since 2011 NASA Goddard Space Flight Center has measured the total hemispheric emissivity of such surfaces from 20 to 300 K using a test apparatus inside a small laboratory cryostat. We report the latest data from these measurements, including Aeroglaze Z307 paint, Black Kapton, and a configuration of painted aluminum honeycomb that was not previously tested. We also present the results of batch-to- batch reproducibility studies in Ball Infrared BlackTM and painted aluminum honeycomb. Finally, we describe a recently-adopted temperature control method which significantly speeds the data acquisition, and we discuss efforts to reduce the noise in future data.

Nuclear Shuttle Logistics Configuration

NASA Technical Reports Server (NTRS)

1971-01-01

This 1971 artist's concept shows the Nuclear Shuttle in both its lunar logistics configuraton and geosynchronous station configuration. As envisioned by Marshall Space Flight Center Program Development persornel, the Nuclear Shuttle would deliver payloads to lunar orbits or other destinations then return to Earth orbit for refueling and additional missions.

The application of some lifting-body reentry concepts to missile design

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1985-01-01

The aerodynamic characteristics of some lifting-body concepts are examined with a view to the applicability of such concepts to the design of missiles. A considerable amount of research has been done in past years with vehicle concepts suitable for manned atmospheric-entry and atmospheric flight. Some of the concepts appear to offer some novel design approaches for missiles for a variety of missions and flight profiles, including long-range orbital/reentry with transatmospheric operation for strategic penetration, low altitude penetration, and battlefield tactical. The concepts considered include right triangular pyramidal configurations, a lenticular configuration, and various 75-degree triangular planform configurations with variations in body camber and control systems. The aerodynamic features are emphasized but some observations are also made relative to other factors such as heat transfer, structures, carriage, observability, propulsion, and volumetric efficiency.

Effect of Free Jet on Refraction and Noise

NASA Technical Reports Server (NTRS)

Khavaran, Abbas; Georgiadis, Nicholas J.; Bridges, James E.; Dippold, Vance F., III

2005-01-01

This article investigates the role of a free jet on the sound radiated from a jet. In particular, the role of an infinite wind tunnel, which simulates the forward flight condition, is compared to that of a finite wind tunnel. The second configuration is usually used in experiments, where the microphones are located in a static ambient medium far outside the free jet. To study the effect of the free jet on noise, both propagation and source strength need to be addressed. In this work, the exact Green's function in a locally parallel flow is derived for a simulated flight case. Numerical examples are presented that show a reduction in the magnitude of the Green's function in the aft arc and an increase in the forward arc for the simulated flight condition. The effect of finite wind tunnel on refraction is sensitive to the source location and is most pronounced in the aft arc. A Reynolds-averaged Navier-Stokes solution (RANS) yields the required mean flow and turbulence scales that are used in the jet mixing noise spectrum calculations. In addition to the sound/flow interaction, the separate effect of source strength and elongation of the noise-generating region of the jet in a forward flight is studied. Comparisons are made with experiments for the static and finite tunnel cases. Finally, the standard free-jet shear corrections that convert the finite wind tunnel measurements to an ideal wind tunnel arrangement are evaluated.

Photogrammetric mapping using unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Graça, N.; Mitishita, E.; Gonçalves, J.

2014-11-01

Nowadays Unmanned Aerial Vehicle (UAV) technology has attracted attention for aerial photogrammetric mapping. The low cost and the feasibility to automatic flight along commanded waypoints can be considered as the main advantages of this technology in photogrammetric applications. Using GNSS/INS technologies the images are taken at the planned position of the exposure station and the exterior orientation parameters (position Xo, Yo, Zo and attitude ω, φ, χ) of images can be direct determined. However, common UAVs (off-the-shelf) do not replace the traditional aircraft platform. Overall, the main shortcomings are related to: difficulties to obtain the authorization to perform the flight in urban and rural areas, platform stability, safety flight, stability of the image block configuration, high number of the images and inaccuracies of the direct determination of the exterior orientation parameters of the images. In this paper are shown the obtained results from the project photogrammetric mapping using aerial images from the SIMEPAR UAV system. The PIPER J3 UAV Hydro aircraft was used. It has a micro pilot MP2128g. The system is fully integrated with 3-axis gyros/accelerometers, GPS, pressure altimeter, pressure airspeed sensors. A Sony Cyber-shot DSC-W300 was calibrated and used to get the image block. The flight height was close to 400 m, resulting GSD near to 0.10 m. The state of the art of the used technology, methodologies and the obtained results are shown and discussed. Finally advantages/shortcomings found in the study and main conclusions are presented

International Space Station (ISS)

NASA Image and Video Library

2001-01-01

This is the STS-102 mission crew insignia. The central image on the crew patch depicts the International Space Station (ISS) in the build configuration that it had at the time of the arrival and docking of Discovery during the STS-102 mission, the first crew exchange flight to the Space Station. The station is shown along the direction of the flight as was seen by the shuttle crew during their final approach and docking, the so-called V-bar approach. The names of the shuttle crew members are depicted in gold around the top of the patch, and surnames of the Expedition crew members being exchanged are shown in the lower barner. The three ribbons swirling up to and around the station signify the rotation of these ISS crew members. The number 2 is for the Expedition 2 crew who flew up to the station, and the number 1 is for the Expedition 1 crew who then returned down to Earth. In conjunction with the face of the Lab module of the Station, these Expedition numbers create the shuttle mission number 102. Shown mated below the ISS is the Italian-built Multipurpose Logistics Module, Leonardo, that flew for the first time on this flight. The flags of the countries that were the major contributors to this effort, the United States, Russia, and Italy are also shown in the lower part of the patch. The build-sequence number of this flight in the overall station assembly sequence, 5A.1, is captured by the constellations in the background.

Air Launch: Examining Performance Potential of Various Configurations and Growth Options

NASA Technical Reports Server (NTRS)

Waters, Eric D.; Creech, Dennis M.; Philips, Alan D.

2013-01-01

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.

Exploratory flow visualization investigation of mast-mounted sights in presence of a rotor

NASA Technical Reports Server (NTRS)

Ghee, Terence A.; Kelley, Henry L.

1995-01-01

A flow visualization investigation with a laser light sheet system was conducted on a 27-percent-scale AH-64 attack helicopter model fitted with two mast-mounted sights in the langley 14- by 22-foot subsonic tunnel. The investigation was conducted to identify aerodynamic phenomena that may have contributed to adverse vibration encountered during full-scale flight of the AH-64D apache/longbow helicopter with an asymmetric mast-mounted sight. Symmetric and asymmetric mast-mounted sights oriented at several skew angles were tested at simulated forward and rearward flight speeds of 30 and 45 knots. A laser light sheet system was used to visualize the flow in planes parallel to and perpendicular to the free-stream flow. Analysis of these flow visualization data identified frequencies of flow patterns in the wake shed from the sight, the streamline angle at the sight, and the location where the shed wake crossed the rotor plane. Differences in wake structure were observed between the sight configurations and various skew angles. Analysis of lateral light sheet plane data implied significant vortex structure in the wake of the asymmetric mast-mounted sight in the configuration that produced maximum in-flight vibration. The data showed no significant vortex structure in the wake of the asymmetric and symmetric configurations that produced no increase in in-flight adverse vibration.

Extended Range Aerial Delivery Using an Unpowered Autonomous Tailless UAV

NASA Astrophysics Data System (ADS)

Kraft, Tyler E.

An alternative approach for precision aerial delivery utilizing a flying wing for controllable forward glide is presented. Although effective, current delivery methods either display a lack of control, or require close standoff distances, potentially endangering aircraft personnel as well as bystanders. Hardware-in-the-loop simulations provide an efficient method for evaluating various wing designs and actuation configurations. Four control surface configurations are presented and evaluated, encompassing traditional aircraft and ram-air parafoil control approaches. Fixed-wing and multirotor unmanned aircraft-based flight tests were conducted to evaluate the controllability and handling performance of the various configurations of both a fixed wing model and a model with collapsing wings. A manufacturing process was developed to allow repeatable results in the field using cheap, mostly disposable materials. A powered flying wing model was used to maximize data collection in later stages of software development. Data collected during flight tests was used to create a model of the system and develop a Nonlinear Dynamic Inversion controller for autonomous flight. The NDI controller was able to provide stable flight in pitch, but will need more development to control yaw, instead an intentional bias was built in to show proof of concept for direct yaw control. The results demonstrate the feasibility of the flying wing-based aerial delivery; however, significant challenges remain regarding the stability and scalability of the system.

Conceptual Design of Low-Boom Aircraft with Flight Trim Requirement

NASA Technical Reports Server (NTRS)

Ordaz, Irian; Geiselhart, Karl A.; Fenbert, James W.

2014-01-01

A new low-boom target generation approach is presented which allows the introduction of a trim requirement during the early conceptual design of supersonic aircraft. The formulation provides an approximation of the center of pressure for a presumed aircraft configuration with a reversed equivalent area matching a low-boom equivalent area target. The center of pressure is approximated from a surrogate lift distribution that is based on the lift component of the classical equivalent area. The assumptions of the formulation are verified to be sufficiently accurate for a supersonic aircraft of high fineness ratio through three case studies. The first two quantify and verify the accuracy and the sensitivity of the surrogate center of pressure corresponding to shape deformation of lifting components. The third verification case shows the capability of the approach to achieve a trim state while maintaining the low-boom characteristics of a previously untrimmed configuration. Finally, the new low-boom target generation approach is demonstrated through the early conceptual design of a demonstrator concept that is low-boom feasible, trimmed, and stable in cruise.

Calculations, and comparison with an ideal minimum, of trimmed drag for conventional and canard configurations having various levels of static stability

NASA Technical Reports Server (NTRS)

Mclaughlin, M. D.

1977-01-01

Classical drag equations were used to calculate total and induced drag and ratios of stabilizer lift to wing lift for a variety of conventional and canard configurations. The Flight efficiencies of such configurations that are trimmed in pitch and have various values of static margin are evaluated. Classical calculation methods are compared with more modern lifting surface theory.

Restructured Freedom configuration characteristics

NASA Technical Reports Server (NTRS)

Troutman, Patrick A.; Heck, Michael L.; Kumar, Renjith R.; Mazanek, Daniel D.

1991-01-01

In Jan. 1991, the LaRc SSFO performed an assessment of the configuration characteristics of the proposed pre-integrated Space Station Freedom (SSF) concept. Of particular concern was the relationship of solar array operation and orientation with respect to spacecraft controllability. For the man-tended configuration (MTC), it was determined that torque equilibrium attitude (TEA) seeking Control Moment Gyroscope (CMG) control laws could not always maintain attitude. The control problems occurred when the solar arrays were tracking the sun to produce full power while flying in an arrow or gravity gradient flight mode. The large solar array articulations that sometimes result from having the functions of the alpha and beta joints reversed on MTC induced large product of inertia changes that can invalidate the control system gains during an orbit. Several modified sun tracking techniques were evaluated with respect to producing a controllable configuration requiring no modifications to the CMG control algorithms. Another assessment involved the permanently manned configuration (PMC) which has a third asymmetric PV unit on one side of the transverse boom. Recommendations include constraining alpha rotations for MTC in the arrow and gravity gradient flight modes and perhaps developing new non-TEA seeking control laws. Recommendations for PMC include raising the operational altitude and moving to a symmetric configuration as soon as possible.

STS-88 Mission Insignia

NASA Technical Reports Server (NTRS)

1998-01-01

Designed by the STS-88 crew members, this patch commemorates the first assembly flight to carry United States-built hardware for constructing the International Space Station (ISS). This flight's primary task was to assemble the cornerstone of the Space Station: the Node with the Functional Cargo Block (FGB). The rising sun symbolizes the dawning of a new era of international cooperation in space and the beginning of a new program: the International Space Station. The Earth scene outlines the countries of the Station Partners: the United States, Russia, those of the European Space Agency (ESA), Japan, and Canada. Along with the Pressurized Mating Adapters (PMA) and the Functional Cargo Block, the Node is shown in the final mated configuration while berthed to the Space Shuttle during the STS-88/2A mission. The Big Dipper Constellation points the way to the North Star, a guiding light for pioneers and explorers for generations. In the words of the crew, These stars symbolize the efforts of everyone, including all the countries involved in the design and construction of the International Space Station, guiding us into the future.

The JWST/NIRSpec instrument: update on status and performances

NASA Astrophysics Data System (ADS)

Birkmann, Stephan M.; Ferruit, Pierre; Rawle, Tim; Sirianni, Marco; Alves de Oliveira, Catarina; Böker, Torsten; Giardino, Giovanna; Lützgendorf, Nora; Marston, Anthony; Stuhlinger, Martin; te Plate, Maurice B. J.; Jensen, Peter; Rumler, Peter; Dorner, Bernhard; Karl, Hermann; Mosner, Peter; Wright, Raymond H.; Rapp, Robert

2016-07-01

The Near-Infrared Spectrograph (NIRSpec) is one of the four instruments on the James Webb Space Telescope (JWST) which is scheduled for launch in 2018. NIRSpec is developed by the European Space Agency (ESA) with Airbus Defense and Space Germany as prime contractor. The instrument offers seven dispersers covering the wavelength range from 0.6 to 5.3 micron with resolutions from R ˜ 100 to R ˜ 2700. NIRSpec will be capable of obtaining spectra for more than 100 objects simultaneously using an array of micro-shutters. It also features an integral field unit with 3" x 3" field of view and a range of slits for high contrast spectroscopy of individual objects and time series observations of e.g. transiting exoplanets. NIRSpec is in its final flight configuration and underwent cryogenic performance testing at the Goddard Space Flight Center in Winter 2015/16 as part of the Integrated Science Instrument Module (ISIM). We present the current status of the instrument and also provide an update on NIRSpec performances based on results from the ISIM level test campaign.

Aerodynamic heating environment definition/thermal protection system selection for the HL-20

NASA Astrophysics Data System (ADS)

Wurster, K. E.; Stone, H. W.

1993-09-01

Definition of the aerothermal environment is critical to any vehicle such as the HL-20 Personnel Launch System that operates within the hypersonic flight regime. Selection of an appropriate thermal protection system design is highly dependent on the accuracy of the heating-environment prediction. It is demonstrated that the entry environment determines the thermal protection system design for this vehicle. The methods used to predict the thermal environment for the HL-20 Personnel Launch System vehicle are described. Comparisons of the engineering solutions with computational fluid dynamic predictions, as well as wind-tunnel test results, show good agreement. The aeroheating predictions over several critical regions of the vehicle, including the stagnation areas of the nose and leading edges, windward centerline and wing surfaces, and leeward surfaces, are discussed. Results of predictions based on the engineering methods found within the MINIVER aerodynamic heating code are used in conjunction with the results of the extensive wind-tunnel tests on this configuration to define a flight thermal environment. Finally, the selection of the thermal protection system based on these predictions and current technology is described.

International Space Station (ISS)

NASA Image and Video Library

1998-11-08

Designed by the STS-88 crew members, this patch commemorates the first assembly flight to carry United States-built hardware for constructing the International Space Station (ISS). This flight's primary task was to assemble the cornerstone of the Space Station: the Node with the Functional Cargo Block (FGB). The rising sun symbolizes the dawning of a new era of international cooperation in space and the beginning of a new program: the International Space Station. The Earth scene outlines the countries of the Station Partners: the United States, Russia, those of the European Space Agency (ESA), Japan, and Canada. Along with the Pressurized Mating Adapters (PMA) and the Functional Cargo Block, the Node is shown in the final mated configuration while berthed to the Space Shuttle during the STS-88/2A mission. The Big Dipper Constellation points the way to the North Star, a guiding light for pioneers and explorers for generations. In the words of the crew, These stars symbolize the efforts of everyone, including all the countries involved in the design and construction of the International Space Station, guiding us into the future.

Simulator study of vortex encounters by a twin-engine, commercial, jet transport airplane

NASA Technical Reports Server (NTRS)

Hastings, E. C., Jr.; Keyser, G. L., Jr.

1982-01-01

A simulator study of vortex encounters was conducted for a twin-engine, commercial, jet transport airplane encountering the vortex flow field of a heavy, four-engine, commercial, jet transport airplane in the final-approach configuration. The encounters were conducted with fixed controls and with a pilot using a state-of-the-art, manual-control system. Piloted encounters with the base-line vortex flow field out of ground effect (unattenuated) resulted in initial bank-angle excursions greater than 40 deg, coupled with initial sideslip-angle excursions greater than 10 deg. The severity of these initial upsets was significantly reduced when the vortex center was moved laterally or vertically away from the flight path of the encountering airplane. Smaller reductions occurred when the flow field was attenuated by the flight spoilers on the generating airplane. The largest reduction in the severity of the initial upsets, however, was from aging in ground effect. The severity of the initial upsets of the following airplane was relatively unaffected by the approach speed. Increasing the lift coefficient of the generating airplane resulted in an increase in the severity of the initial upsets.

Space Station redesign option A: Modular buildup concept

NASA Technical Reports Server (NTRS)

1993-01-01

In early 1993, President Clinton mandated that NASA look at lower cost alternatives to Space Station Freedom. He also established an independent advisory committee - the Blue Ribbon Panel - to review the redesign work and evaluate alternatives. Daniel Goldin, NASA Administrator, established a Station Redesign Team that began operating in late March from Crystal City, Virginia. NASA intercenter teams - one each at Marshall Space Flight Center, Johnson Space Center, and Langley Research Center provided engineering and other support. The results of the Option A study done at Marshall Space Flight Center are summarized. Two configurations (A-1 and A-2) are covered. Additional data is provided in the briefing package MSFC SRT-001, Final System Review to SRT-002, Space Station Option A Modular Buildup Concept, Volumes 1-5, Revision B, June 10, 1993. In June 1993, President Clinton decided to proceed with a modular concept consistent with Option A, and asked NASA to provide an Implementation Plan by September. All data from the Option A redesign activity was provided to NASA's Transition Team for use in developing the Implementation Plan.

Aerothermodynamic Testing and Boundary Layer Trip Sizing of the HIFiRE Flight 1 Vehicle

NASA Technical Reports Server (NTRS)

Berger, Karen T.; Greene, Frank A.; Kimmel, Roger; Alba, Christopher; Johnson, Heath

2008-01-01

An experimental wind tunnel test was conducted in the NASA Langley Research Center s 20-Inch Mach 6 Air Tunnel in support of the Hypersonic International Flight Research Experimentation Program. The information in this article is focused on the Flight 1 configuration, the first in a series of flight experiments. The article documents experimental measurements made over a Reynolds numbers range of 2.1x10(exp 6)/ft to 5.6x10(exp 6)/ft and angles of attack of -5 to +5 deg on several scaled ceramic heat transfer models of the Flight 1 configuration. Global heat transfer was measured using phosphor thermography and the resulting images and heat transfer distributions were used to infer the state of the boundary layer on the vehicle windside and leeside surfaces. Boundary layer trips were used to force the boundary layer turbulent and the experimental data highlighted in this article were used to size and place the boundary layer trip for the flight vehicle. The required height of the flight boundary layer trip was determined to be 0.079 in and the trip was moved from the design location of 7.87 in to 20.47 in to ensure that augmented heating would not impact the laminar side of the vehicle. Allowable roughness was selected to be 3.2x10(exp -3) in.

The Development and Flight Testing of an Aerially Deployed Unmanned Aerial System

NASA Astrophysics Data System (ADS)

Smith, Andrew

An investigation into the feasibility of aerial deployed unmanned aerial vehicles was completed. The investigation included the development and flight testing of multiple unmanned aerial systems to investigate the different components of potential aerial deployment missions. The project consisted of two main objectives; the first objective dealt with the development of an airframe capable of surviving aerial deployment from a rocket and then self assembling from its stowed configuration into its flight configuration. The second objective focused on the development of an autopilot capable of performing basic guidance, navigation, and control following aerial deployment. To accomplish these two objectives multiple airframes were developed to verify their completion experimentally. The first portion of the project, investigating the feasibility of surviving an aerial deployment, was completed using a fixed wing glider that following a successful deployment had 52 seconds of controlled flight. Before developing the autopilot in the second phase of the project, the glider was significantly upgraded to fix faults discovered in the glider flight testing and to enhance the system capabilities. Unfortunately to conform to outdoor flight restrictions imposed by the university and the Federal Aviation Administration it was required to switch airframes before flight testing of the new fixed wing platform could begin. As a result, an autopilot was developed for a quadrotor and verified experimentally completely indoors to remain within the limits of governing policies.

Flight Measurements of the Flying Qualities of a Lockheed P-80A Airplane (Army No. 44-85099) - Stalling Characteristics

NASA Technical Reports Server (NTRS)

Anderson, Seth B.; Cooper, George E.

1947-01-01

This report contains the flight-test results of the stalling characteristics measured during the flying-qualities investigation of the Lockheed P-8OA airplane (Army No. 44-85099). The tests were conducted in straight and turning flight with and without wing-tip tanks. These tests showed satisfactory stalling characteristics and adequate stall warning for all configurations and conditions tested.

Flight-determined stability and control characteristics of the M2-F3 lifting body vehicle

NASA Technical Reports Server (NTRS)

Sim, A. G.

1973-01-01

Flight data were obtained over a Mach number range from 0.4 to 1.55 and an angle-of-attack range from -2 deg to 16 deg. Lateral-directional and longitudinal derivatives, reaction control rocket effectiveness, and longitudinal trim information obtained from flight data and wind-tunnel predictions are compared. The effects of power, configuration change, and speed brake are discussed.

Adaptive structures flight experiments

NASA Astrophysics Data System (ADS)

Martin, Maurice

The topics are presented in viewgraph form and include the following: adaptive structures flight experiments; enhanced resolution using active vibration suppression; Advanced Controls Technology Experiment (ACTEX); ACTEX program status; ACTEX-2; ACTEX-2 program status; modular control patch; STRV-1b Cryocooler Vibration Suppression Experiment; STRV-1b program status; Precision Optical Bench Experiment (PROBE); Clementine Spacecraft Configuration; TECHSAT all-composite spacecraft; Inexpensive Structures and Materials Flight Experiment (INFLEX); and INFLEX program status.

Adaptive Structures Flight Experiments

NASA Technical Reports Server (NTRS)

Martin, Maurice

1992-01-01

The topics are presented in viewgraph form and include the following: adaptive structures flight experiments; enhanced resolution using active vibration suppression; Advanced Controls Technology Experiment (ACTEX); ACTEX program status; ACTEX-2; ACTEX-2 program status; modular control patch; STRV-1b Cryocooler Vibration Suppression Experiment; STRV-1b program status; Precision Optical Bench Experiment (PROBE); Clementine Spacecraft Configuration; TECHSAT all-composite spacecraft; Inexpensive Structures and Materials Flight Experiment (INFLEX); and INFLEX program status.

Multi-Exciter Vibroacoustic Simulation of Hypersonic Flight Vibration

DOE Office of Scientific and Technical Information (OSTI.GOV)

GREGORY,DANNY LYNN; CAP,JEROME S.; TOGAMI,THOMAS C.

1999-11-11

Many aerospace structures must survive severe high frequency, hypersonic, random vibration during their flights. The random vibrations are generated by the turbulent boundary layer developed along the exterior of the structures during flight. These environments have not been simulated very well in the past using a fixed-based, single exciter input with an upper frequency range of 2 kHz. This study investigates the possibility of using acoustic ardor independently controlled multiple exciters to more accurately simulate hypersonic flight vibration. The test configuration, equipment, and methodology are described. Comparisons with actual flight measurements and previous single exciter simulations are also presented.

Can a glass cockpit display help (or hinder) performance of novices in simulated flight training?

PubMed

Wright, Stephen; O'Hare, David

2015-03-01

The analog dials in traditional GA aircraft cockpits are being replaced by integrated electronic displays, commonly referred to as glass cockpits. Pilots may be trained on glass cockpit aircraft or encounter them after training on traditional displays. The effects of glass cockpit displays on initial performance and potential transfer effects between cockpit display configurations have yet to be adequately investigated. Flight-naïve participants were trained on either a simulated traditional display cockpit or a simulated glass display cockpit. Flight performance was measured in a test flight using either the same or different cockpit display. Loss of control events and accuracy in controlling altitude, airspeed and heading, workload, and situational awareness were assessed. Preferences for cockpit display configurations and opinions on ease of use were also measured. The results revealed consistently poorer performance on the test flight for participants using the glass cockpit compared to the traditional cockpit. In contrast the post-flight questionnaire data revealed a strong subjective preference for the glass cockpit over the traditional cockpit displays. There was only a weak effect of prior training. The specific glass cockpit display used in this study was subjectively appealing but yielded poorer flight performance in participants with no previous flight experience than a traditional display. Performance data can contradict opinion data. The design of glass cockpit displays may present some difficulties for pilots in the very early stages of training. Copyright © 2014 Elsevier Ltd and The Ergonomics Society. All rights reserved.

14 CFR 25.251 - Vibration and buffeting.

Code of Federal Regulations, 2011 CFR

2011-01-01

... cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the...) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any... the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the...

14 CFR 25.251 - Vibration and buffeting.

Code of Federal Regulations, 2014 CFR

2014-01-01

... cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the...) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any... the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the...

14 CFR 25.251 - Vibration and buffeting.

Code of Federal Regulations, 2010 CFR

2010-01-01

... cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the...) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any... the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the...

14 CFR 25.251 - Vibration and buffeting.

Code of Federal Regulations, 2012 CFR

2012-01-01

... cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the...) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any... the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the...

14 CFR 25.251 - Vibration and buffeting.

Code of Federal Regulations, 2013 CFR

2013-01-01

... cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the...) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any... the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the...

Optimization of a Lunar Pallet Lander Reinforcement Structure Using a Genetic Algorithm

NASA Technical Reports Server (NTRS)

Burt, Adam O.; Hull, Patrick V.

2014-01-01

This paper presents a design automation process using optimization via a genetic algorithm to design the conceptual structure of a Lunar Pallet Lander. The goal is to determine a design that will have the primary natural frequencies at or above a target value as well as minimize the total mass. Several iterations of the process are presented. First, a concept optimization is performed to determine what class of structure would produce suitable candidate designs. From this a stiffened sheet metal approach was selected leading to optimization of beam placement through generating a two-dimensional mesh and varying the physical location of reinforcing beams. Finally, the design space is reformulated as a binary problem using 1-dimensional beam elements to truncate the design space to allow faster convergence and additional mechanical failure criteria to be included in the optimization responses. Results are presented for each design space configuration. The final flight design was derived from these results.

14 CFR 36.1501 - Procedures, noise levels and other information.

Code of Federal Regulations, 2011 CFR

2011-01-01

... airplane (rotorcraft) flight manual. (b) Where supplemental test data are approved for modification or extension of an existing flight data base, such as acoustic data from engine static tests used in the..., weights, configurations, and other information or data employed for obtaining the certified noise levels...

14 CFR 36.1501 - Procedures, noise levels and other information.

Code of Federal Regulations, 2014 CFR

2014-01-01

... airplane (rotorcraft) flight manual. (b) Where supplemental test data are approved for modification or extension of an existing flight data base, such as acoustic data from engine static tests used in the..., weights, configurations, and other information or data employed for obtaining the certified noise levels...

14 CFR 36.1501 - Procedures, noise levels and other information.

Code of Federal Regulations, 2012 CFR

2012-01-01

... airplane (rotorcraft) flight manual. (b) Where supplemental test data are approved for modification or extension of an existing flight data base, such as acoustic data from engine static tests used in the..., weights, configurations, and other information or data employed for obtaining the certified noise levels...

14 CFR 36.1501 - Procedures, noise levels and other information.

Code of Federal Regulations, 2010 CFR

2010-01-01

... airplane (rotorcraft) flight manual. (b) Where supplemental test data are approved for modification or extension of an existing flight data base, such as acoustic data from engine static tests used in the..., weights, configurations, and other information or data employed for obtaining the certified noise levels...

14 CFR 36.1501 - Procedures, noise levels and other information.

Code of Federal Regulations, 2013 CFR

2013-01-01

... airplane (rotorcraft) flight manual. (b) Where supplemental test data are approved for modification or extension of an existing flight data base, such as acoustic data from engine static tests used in the..., weights, configurations, and other information or data employed for obtaining the certified noise levels...

Flight Experiment Demonstration System (FEDS) analysis report

NASA Technical Reports Server (NTRS)

Shank, D. E.

1986-01-01

The purpose of the Flight Experiment Demonstration System (FEDS) was to show, in a simulated spacecraft environment, the feasibility of using a microprocessor to automate the onboard orbit determination functions. The software and hardware configuration used to support FEDS during the demonstration and the results of the demonstration are discussed.

Topology optimization of pressure adaptive honeycomb for a morphing flap

NASA Astrophysics Data System (ADS)

Vos, Roelof; Scheepstra, Jan; Barrett, Ron

2011-03-01

The paper begins with a brief historical overview of pressure adaptive materials and structures. By examining avian anatomy, it is seen that pressure-adaptive structures have been used successfully in the Natural world to hold structural positions for extended periods of time and yet allow for dynamic shape changes from one flight state to the next. More modern pneumatic actuators, including FAA certified autopilot servoactuators are frequently used by aircraft around the world. Pneumatic artificial muscles (PAM) show good promise as aircraft actuators, but follow the traditional model of load concentration and distribution commonly found in aircraft. A new system is proposed which leaves distributed loads distributed and manipulates structures through a distributed actuator. By using Pressure Adaptive Honeycomb (PAH), it is shown that large structural deformations in excess of 50% strains can be achieved while maintaining full structural integrity and enabling secondary flight control mechanisms like flaps. The successful implementation of pressure-adaptive honeycomb in the trailing edge of a wing section sparked the motivation for subsequent research into the optimal topology of the pressure adaptive honeycomb within the trailing edge of a morphing flap. As an input for the optimization two known shapes are required: a desired shape in cruise configuration and a desired shape in landing configuration. In addition, the boundary conditions and load cases (including aerodynamic loads and internal pressure loads) should be specified for each condition. Finally, a set of six design variables is specified relating to the honeycomb and upper skin topology of the morphing flap. A finite-element model of the pressure-adaptive honeycomb structure is developed specifically tailored to generate fast but reliable results for a given combination of external loading, input variables, and boundary conditions. Based on two bench tests it is shown that this model correlates well to experimental results. The optimization process finds the skin and honeycomb topology that minimizes the error between the acquired shape and the desired shape in each configuration.

Benefits of Hybrid-Electric Propulsion to Achieve 4x Increase in Cruise Efficiency for a VTOL Aircraft

NASA Technical Reports Server (NTRS)

Fredericks, William J.; Moore, Mark D.; Busan, Ronald C.

2013-01-01

Electric propulsion enables radical new vehicle concepts, particularly for Vertical Takeoff and Landing (VTOL) aircraft because of their significant mismatch between takeoff and cruise power conditions. However, electric propulsion does not merely provide the ability to normalize the power required across the phases of flight, in the way that automobiles also use hybrid electric technologies. The ability to distribute the thrust across the airframe, without mechanical complexity and with a scale-free propulsion system, is a new degree of freedom for aircraft designers. Electric propulsion is scale-free in terms of being able to achieve highly similar levels of motor power to weight and efficiency across a dramatic scaling range. Applying these combined principles of electric propulsion across a VTOL aircraft permits an improvement in aerodynamic efficiency that is approximately four times the state of the art of conventional helicopter configurations. Helicopters typically achieve a lift to drag ratio (L/D) of between 4 and 5, while the VTOL aircraft designed and developed in this research were designed to achieve an L/D of approximately 20. Fundamentally, the ability to eliminate the problem of advancing and retreating rotor blades is shown, without resorting to unacceptable prior solutions such as tail-sitters. This combination of concept and technology also enables a four times increase in range and endurance while maintaining the full VTOL and hover capability provided by a helicopter. Also important is the ability to achieve low disc-loading for low ground impingement velocities, low noise and hover power minimization (thus reducing energy consumption in VTOL phases). This combination of low noise and electric propulsion (i.e. zero emissions) will produce a much more community-friendly class of vehicles. This research provides a review of the concept brainstorming, configuration aerodynamic and mission analysis, as well as subscale prototype construction and flight testing that verifies transition flight control. A final down-selected vehicle is also presented.

Design and Flight Tests of an Adaptive Control System Employing Normal-Acceleration Command

NASA Technical Reports Server (NTRS)

McNeill, Water E.; McLean, John D.; Hegarty, Daniel M.; Heinle, Donovan R.

1961-01-01

An adaptive control system employing normal-acceleration command has been designed with the aid of an analog computer and has been flight tested. The design of the system was based on the concept of using a mathematical model in combination with a high gain and a limiter. The study was undertaken to investigate the application of a system of this type to the task of maintaining nearly constant dynamic longitudinal response of a piloted airplane over the flight envelope without relying on air data measurements for gain adjustment. The range of flight conditions investigated was between Mach numbers of 0.36 and 1.15 and altitudes of 10,000 and 40,000 feet. The final adaptive system configuration was derived from analog computer tests, in which the physical airplane control system and much of the control circuitry were included in the loop. The method employed to generate the feedback signals resulted in a model whose characteristics varied somewhat with changes in flight condition. Flight results showed that the system limited the variation in longitudinal natural frequency of the adaptive airplane to about half that of the basic airplane and that, for the subsonic cases, the damping ratio was maintained between 0.56 and 0.69. The system also automatically compensated for the transonic trim change. Objectionable features of the system were an exaggerated sensitivity of pitch attitude to gust disturbances, abnormally large pitch attitude response for a given pilot input at low speeds, and an initial delay in normal-acceleration response to pilot control at all flight conditions. The adaptive system chatter of +/-0.05 to +/-0.10 of elevon at about 9 cycles per second (resulting in a maximum airplane normal-acceleration response of from +/-0.025 g to +/- 0.035 g) was considered by the pilots to be mildly objectionable but tolerable.

Feasibility of a Networked Air Traffic Infrastructure Validation Environment for Advanced NextGen Concepts

NASA Technical Reports Server (NTRS)

McCormack, Michael J.; Gibson, Alec K.; Dennis, Noah E.; Underwood, Matthew C.; Miller,Lana B.; Ballin, Mark G.

2013-01-01

Abstract-Next Generation Air Transportation System (NextGen) applications reliant upon aircraft data links such as Automatic Dependent Surveillance-Broadcast (ADS-B) offer a sweeping modernization of the National Airspace System (NAS), but the aviation stakeholder community has not yet established a positive business case for equipage and message content standards remain in flux. It is necessary to transition promising Air Traffic Management (ATM) Concepts of Operations (ConOps) from simulation environments to full-scale flight tests in order to validate user benefits and solidify message standards. However, flight tests are prohibitively expensive and message standards for Commercial-off-the-Shelf (COTS) systems cannot support many advanced ConOps. It is therefore proposed to simulate future aircraft surveillance and communications equipage and employ an existing commercial data link to exchange data during dedicated flight tests. This capability, referred to as the Networked Air Traffic Infrastructure Validation Environment (NATIVE), would emulate aircraft data links such as ADS-B using in-flight Internet and easily-installed test equipment. By utilizing low-cost equipment that is easy to install and certify for testing, advanced ATM ConOps can be validated, message content standards can be solidified, and new standards can be established through full-scale flight trials without necessary or expensive equipage or extensive flight test preparation. This paper presents results of a feasibility study of the NATIVE concept. To determine requirements, six NATIVE design configurations were developed for two NASA ConOps that rely on ADS-B. The performance characteristics of three existing in-flight Internet services were investigated to determine whether performance is adequate to support the concept. Next, a study of requisite hardware and software was conducted to examine whether and how the NATIVE concept might be realized. Finally, to determine a business case, economic factors were evaluated and a preliminary cost-benefit analysis was performed.

An in-flight simulator investigation of roll and yaw control power requirements for STOL approach and landing: Development of capability and preliminary results

NASA Technical Reports Server (NTRS)

Ellis, D. R.; Raisinghani, S. C.

1979-01-01

A six-degree-of-freedom variable-response research aircraft was used to determine the minimum lateral-directional control power required for desirable and acceptable levels of handling qualities for the STOL landing approach task in a variety of simulated atmospheric disturbance conditions for a range of lateral-directional response characteristics. Topics covered include the in-flight simulator, crosswind simulation, turbulence simulation, test configurations, and evaluation procedures. Conclusions based on a limited sampling of simulated STOL transport configurations flown to touchdown out of 6 deg, 75 kt MLS approaches, usually with a sidestep maneuver are discussed.

A new innovative instrument for space plasma instrumentation

NASA Technical Reports Server (NTRS)

Torbert, Roy B.

1993-01-01

The Faraday Ring Ammeter was the subject of this grant for a new innovative instrument for space plasma instrumentation. This report summarizes our progress in this work. Briefly, we have conducted an intensive series of experiments and trials over three years, testing some five configurations of the instrument to measure currents, resulting in two Ph.D. theses, supported by this grant, and two flight configurations of the instrument. The first flight would have been on a NASA-Air Force collaborative sounding rocket, but was not flown because of instrumental difficulties. The second has been successfully integrated on the NASA Auroral Turbulence payload which is to be launched in February, 1994.

Maneuver Acoustic Flight Test of the Bell 430 Helicopter

NASA Technical Reports Server (NTRS)

Watts, Michael E.; Snider, Royce; Greenwood, Eric; Baden, Joel

2012-01-01

A cooperative flight test by NASA, Bell Helicopter and the U.S. Army to characterize the steady state acoustics and measure the maneuver noise of a Bell Helicopter 430 aircraft was accomplished. The test occurred during June/July, 2011 at Eglin Air Force Base, Florida. This test gathered a total of 410 data points over 10 test days and compiled an extensive data base of dynamic maneuver measurements. Three microphone configurations with up to 31 microphones in each configuration were used to acquire acoustic data. Aircraft data included DGPS, aircraft state and rotor state information. This paper provides an overview of the test.

The transition of a real-time single-rotor helicopter simulation program to a supercomputer

NASA Technical Reports Server (NTRS)

Martinez, Debbie

1995-01-01

This report presents the conversion effort and results of a real-time flight simulation application transition to a CONVEX supercomputer. Enclosed is a detailed description of the conversion process and a brief description of the Langley Research Center's (LaRC) flight simulation application program structure. Currently, this simulation program may be configured to represent Sikorsky S-61 helicopter (a five-blade, single-rotor, commercial passenger-type helicopter) or an Army Cobra helicopter (either the AH-1 G or AH-1 S model). This report refers to the Sikorsky S-61 simulation program since it is the most frequently used configuration.

Effects of acoustic treatment on the interior noise levels of a twin-engine propeller aircraft - Experimental flight results and theoretical predictions

NASA Technical Reports Server (NTRS)

Beyer, T. B.; Powell, C. A.; Daniels, E. F.; Pope, L. D.

1984-01-01

In-flight noise level measurements were made within two cabin configurations of a general aviation business aircraft. The Fairchild Merlin IVC twin-engine aircraft was tested with bare walls and fiberglass insulation and in an executive trim configuration. Narrow-band and octave format data were subjected to analyses which permitted identification of the blade passage harmonics (BPH). Cabin noise level reductions (insertion losses) due to added insulation varied with position in the cabin, the BPH number, cabin pressure, and engine torque. The measurements were closely predicted using the propeller aircraft interior noise (PAIN) mode.

SSME/side loads analysis for flight configuration, revision A. [structural analysis of space shuttle main engine under side load excitation

NASA Technical Reports Server (NTRS)

Holland, W.

1974-01-01

This document describes the dynamic loads analysis accomplished for the Space Shuttle Main Engine (SSME) considering the side load excitation associated with transient flow separation on the engine bell during ground ignition. The results contained herein pertain only to the flight configuration. A Monte Carlo procedure was employed to select the input variables describing the side load excitation and the loads were statistically combined. This revision includes an active thrust vector control system representation and updated orbiter thrust structure stiffness characteristics. No future revisions are planned but may be necessary as system definition and input parameters change.

The aerodynamics of some guided projectiles

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1984-01-01

Some characteristic projectile shapes are considered with various added components intended to provide lift, stability, and control. The intent of the additions is to provide some means for altering the normal ballistic flight path of a projectile for various purposes such as: achieving greater accuracy at the impact point, selecting alternate impact points, extending range, improved evasion, and so on. The configurations presented illustrate the effects of a flare, wings, and tails for providing stability and lift, and the effects of aft-tails, a close-coupled flap, and all-moving forward wings for control. The relative merits of the various configurations, all of which provided for flight path alterations are discussed.

The aerodynamics of some guided projectiles

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1984-01-01

Some characteristic projectile shapes are considered with various added components intended to provide lift, stability, and control. The intent of the additions is to provide some means for altering the normal ballistic flight path of a projectile for various purposes such as: achieving greater accuracy at the impact point, selecting alternate impact points, extending range, improved evasion, and so on. The configurations presented illustrate the effects of a flare, wings, and tails for providing stability and lift, and the effects of aft-tails, a close-coupled flap, and all-moving forward wings for control. The relative merits of the various configurations, all of which provided for flight path alterations, are discussed.

KC-135A Winglet Flight Flutter Program

NASA Technical Reports Server (NTRS)

Kehoe, M. W.

1982-01-01

The evaluation techniques, results and conclusions for the flight flutter testing conducted on a KC-135A airplane configured with and without winglets are discussed. Test results are presented for the critical symmetric and antisymmetric modes for a fuel distribution that consisted of 10,000 pounds in each wing main tank and empty reserve tanks. The results indicated that a lightly damped oscillation was experienced for a winglet configuration of a 0 deg cant and -4 deg incidence. The effects of cant and incidence angle variation on the critical modes are also discussed. Lightly damped oscillations were not encountered for any other winglet cant and incidence angles tested.

Detached Eddy Simulation for the F-16XL Aircraft Configuration

NASA Technical Reports Server (NTRS)

Elmiligui, Alaa; Abdol-Hamid, Khaled; Parlette, Edward B.

2015-01-01

Numerical simulations for the flow around the F-16XL configuration as a contribution to the Cranked Arrow Wing Aerodynamic Project International 2 (CAWAPI-2) have been performed. The NASA Langley Tetrahedral Unstructured Software System (TetrUSS) with its USM3D solver was used to perform the unsteady flow field simulations for the subsonic high angle-of-attack case corresponding to flight condition (FC) 25. Two approaches were utilized to capture the unsteady vortex flow over the wing of the F-16XL. The first approach was to use Unsteady Reynolds-Averaged Navier-Stokes (URANS) coupled with standard turbulence closure models. The second approach was to use Detached Eddy Simulation (DES), which creates a hybrid model that attempts to combine the most favorable elements of URANS models and Large Eddy Simulation (LES). Computed surface static pressure profiles are presented and compared with flight data. Time-averaged and instantaneous results obtained on coarse, medium and fine grids are compared with the flight data. The intent of this study is to demonstrate that the DES module within the USM3D solver can be used to provide valuable data in predicting vortex-flow physics on a complex configuration.

Orion Launch Abort Vehicle Attitude Control Motor Testing

NASA Technical Reports Server (NTRS)

Murphy, Kelly J.; Brauckmann, Gregory J.; Paschal, Keith B.; Chan, David T.; Walker, Eric L.; Foley, Robert; Mayfield, David; Cross, Jared

2011-01-01

Current Orion Launch Abort Vehicle (LAV) configurations use an eight-jet, solid-fueled Attitude Control Motor (ACM) to provide required vehicle control for all proposed abort trajectories. Due to the forward position of the ACM on the LAV, it is necessary to assess the effects of jet-interactions (JI) between the various ACM nozzle plumes and the external flow along the outside surfaces of the vehicle. These JI-induced changes in flight control characteristics must be accounted for in developing ACM operations and LAV flight characteristics. A test program to generate jet interaction aerodynamic increment data for multiple LAV configurations was conducted in the NASA Ames and NASA Langley Unitary Plan Wind Tunnels from August 2007 through December 2009. Using cold air as the simulant gas, powered subscale models were used to generate interaction data at subsonic, transonic, and supersonic test conditions. This paper presents an overview of the complete ACM JI experimental test program for Orion LAV configurations, highlighting ACM system modeling, nozzle scaling assumptions, experimental test techniques, and data reduction methodologies. Lessons learned are discussed, and sample jet interaction data are shown. These data, in conjunction with computational predictions, were used to create the ACM JI increments for all relevant flight databases.

Shielding of Turbomachinery Broadband Noise from a Hybrid Wing Body Aircraft Configuration

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, Thomas F.; Burley, Casey L.; Bahr, Christopher J.; Stead, Daniel J.; Pope, D. Stuart

2014-01-01

The results of an experimental study on the effects of engine placement and vertical tail configuration on shielding of exhaust broadband noise radiation are presented. This study is part of the high fidelity aeroacoustic test of a 5.8% scale Hybrid Wing Body (HWB) aircraft configuration performed in the 14- by 22-Foot Subsonic Tunnel at NASA Langley Research Center. Broadband Engine Noise Simulators (BENS) were used to determine insertion loss due to shielding by the HWB airframe of the broadband component of turbomachinery noise for different airframe configurations and flight conditions. Acoustics data were obtained from flyover and sideline microphones traversed to predefined streamwise stations. Noise measurements performed for different engine locations clearly show the noise benefit associated with positioning the engine nacelles further upstream on the HWB centerbody. Positioning the engine exhaust 2.5 nozzle diameters upstream (compared to 0.5 nozzle diameters downstream) of the HWB trailing edge was found of particular benefit in this study. Analysis of the shielding performance obtained with and without tunnel flow show that the effectiveness of the fuselage shielding of the exhaust noise, although still significant, is greatly reduced by the presence of the free stream flow compared to static conditions. This loss of shielding is due to the turbulence in the model near-wake/boundary layer flow. A comparison of shielding obtained with alternate vertical tail configurations shows limited differences in level; nevertheless, overall trends regarding the effect of cant angle and vertical location are revealed. Finally, it is shown that the vertical tails provide a clear shielding benefit towards the sideline while causing a slight increase in noise below the aircraft.

Atmosphere, Magnetosphere and Plasmas in Space (AMPS). Space payload definition study. Volume 2: Mission support requirements document

NASA Technical Reports Server (NTRS)

1976-01-01

The flight payload, its operation, and the support required from the Space Transporatation System (STS) is defined including the flight objectives and requirements, the experiment operations, and the payload configurations. The support required from the STS includes the accommodation of the payload by the orbiter/Spacelab, use of the flight operations network and ground facilities, and the use of the launch site facilities.

Range Safety for an Autonomous Flight Safety System

NASA Technical Reports Server (NTRS)

Lanzi, Raymond J.; Simpson, James C.

2010-01-01

The Range Safety Algorithm software encapsulates the various constructs and algorithms required to accomplish Time Space Position Information (TSPI) data management from multiple tracking sources, autonomous mission mode detection and management, and flight-termination mission rule evaluation. The software evaluates various user-configurable rule sets that govern the qualification of TSPI data sources, provides a prelaunch autonomous hold-launch function, performs the flight-monitoring-and-termination functions, and performs end-of-mission safing

Automated Test for NASA CFS

NASA Technical Reports Server (NTRS)

McComas, David C.; Strege, Susanne L.; Carpenter, Paul B. Hartman, Randy

2015-01-01

The core Flight System (cFS) is a flight software (FSW) product line developed by the Flight Software Systems Branch (FSSB) at NASA's Goddard Space Flight Center (GSFC). The cFS uses compile-time configuration parameters to implement variable requirements to enable portability across embedded computing platforms and to implement different end-user functional needs. The verification and validation of these requirements is proving to be a significant challenge. This paper describes the challenges facing the cFS and the results of a pilot effort to apply EXB Solution's testing approach to the cFS applications.

Integrated digital/electric aircraft concepts study

NASA Technical Reports Server (NTRS)

Cronin, M. J.; Hays, A. P.; Green, F. B.; Radovcich, N. A.; Helsley, C. W.; Rutchik, W. L.

1985-01-01

The integrated digital/electrical aircraft (IDEA) is an aircraft concept which employs all electric secondary power systems and advanced digital flight control systems. After trade analysis, preferred systems were applied to the baseline configuration. An additional configuration, the alternate IDEA, was also considered. For this concept the design ground rules were relaxed in order to quantify additional synergistic benefits. It was proposed that an IDEA configuration and technical risks associated with the IDEA systems concepts be defined and the research and development required activities to reduce these risks be identified. The selected subsystems include: power generation, power distribution, actuators, environmental control system and flight controls systems. When the aircraft was resized, block fuel was predicted to decrease by 11.3 percent, with 7.9 percent decrease in direct operating cost. The alternate IDEA shows a further 3.4 percent reduction in block fuel and 3.1 percent reduction in direct operating cost.

Display/control requirements for automated VTOL aircraft

NASA Technical Reports Server (NTRS)

Hoffman, W. C.; Kleinman, D. L.; Young, L. R.

1976-01-01

A systematic design methodology for pilot displays in advanced commercial VTOL aircraft was developed and refined. The analyst is provided with a step-by-step procedure for conducting conceptual display/control configurations evaluations for simultaneous monitoring and control pilot tasks. The approach consists of three phases: formulation of information requirements, configuration evaluation, and system selection. Both the monitoring and control performance models are based upon the optimal control model of the human operator. Extensions to the conventional optimal control model required in the display design methodology include explicit optimization of control/monitoring attention; simultaneous monitoring and control performance predictions; and indifference threshold effects. The methodology was applied to NASA's experimental CH-47 helicopter in support of the VALT program. The CH-47 application examined the system performance of six flight conditions. Four candidate configurations are suggested for evaluation in pilot-in-the-loop simulations and eventual flight tests.

Application of frequency domain handling qualities criteria to the longitudinal landing task

NASA Technical Reports Server (NTRS)

Sarrafian, S. K.; Powers, B. G.

1985-01-01

Under NASA sponsorship, an in-flight simulation of the longitudinal handling qualities of several configurations for the approach and landing tasks was performed on the USAF/AFWAL Total In-Flight Simulator by the Calspan Corporation. The basic configuration was a generic transport airplane with static instability. The control laws included proportional plus integral gain loops to produce pitch-rate and angle-of-attack feedback loops. The evaluation task was a conventional visual approach to a flared touchdown at a designated spot on the runway with a lateral offset. The general conclusions were that the existing criteria are based on pitch-attitude response and that these characteristics do not adequately discriminate between the good and bad configurations of this study. This paper describes the work that has been done to further develop frequency-based criteria in an effort to provide better correlation with the observed data.

Artist concept of Shuttle Solar Backscatter UV (SSBUV) flight configuration

NASA Technical Reports Server (NTRS)

1989-01-01

Artist concept of STS-34 payload bay (PLB) experiment is titled SSBUV FLIGHT CONFIGURATION. The labeled drawing of the Shuttle Solar Backscatter Ultraviolet (UV) (SSBUV) get away special (GAS) canisters identifies the adapter beam, motorized door mechanism, instrument canister, support canister, bottom hat, and interconnect cable. The GAS canisters will be mounted on the starboard wall of Atlantis', Orbiter Vehicle (OV) 104's, PLB. One canister contains an instrument nearly identical to that flown on the satellite. The second canister provides power, data, and command systems. During STS-34, SSBUV instrument will calibrate similar ozone measuring space-based instruments on the National Oceanic and Atmospheric Administration's (NOAA's) TIROS satellites (NOAA-9 and NOAA-11). SSBUV uses the Space Shuttle's orbital flight path to assess instrument performance by directly comparing data from identical instruments aboard TIROS spacecraft, as the Shuttle and the satellite pass over the same E

Bioinspired morphing wings for extended flight envelope and roll control of small drones.

PubMed

Di Luca, M; Mintchev, S; Heitz, G; Noca, F; Floreano, D

2017-02-06

Small-winged drones can face highly varied aerodynamic requirements, such as high manoeuvrability for flight among obstacles and high wind resistance for constant ground speed against strong headwinds that cannot all be optimally addressed by a single aerodynamic profile. Several bird species solve this problem by changing the shape of their wings to adapt to the different aerodynamic requirements. Here, we describe a novel morphing wing design composed of artificial feathers that can rapidly modify its geometry to fulfil different aerodynamic requirements. We show that a fully deployed configuration enhances manoeuvrability while a folded configuration offers low drag at high speeds and is beneficial in strong headwinds. We also show that asymmetric folding of the wings can be used for roll control of the drone. The aerodynamic performance of the morphing wing is characterized in simulations, in wind tunnel measurements and validated in outdoor flights with a small drone.

A free flight investigation of transonic sting interference

NASA Technical Reports Server (NTRS)

Jaffe, P.

1975-01-01

Transonic sting interference has been studied in a supersonic wind tunnel to obtain free flight and sting support data on identical models. The two principal configurations, representing fuselage bodies, were cigar shaped with tail fins. The others were a sharp 10-deg cone, a sphere, and a blunt entry body. Comparative data indicated that the sting had an appreciable effect on drag for the fuselage-like configurations; drag rise occurred 0.02 Mach number earlier in free flight, and drag level was 15% greater. The spheres and the blunt bodies were insensitive to the presence of stings regardless of their size. The 10-deg cones were in between, experiencing no drag difference with a minimum diameter sting, but a moderate difference with the largest diameter sting tested. All data tend to confirm the notion that for the more slender bodies the sting not only affects flow but the forebody flow as well.

Porous tube plant nutrient delivery system development: A device for nutrient delivery in microgravity

NASA Technical Reports Server (NTRS)

Dreschel, T. W.; Brown, C. S.; Piastuch, W. C.; Hinkle, C. R.; Knott, W. M.

1994-01-01

The Porous Tube Plant Nutrient Delivery Systems or PTPNDS (U.S. Patent #4,926,585) has been under development for the past six years with the goal of providing a means for culturing plants in microgravity, specifically providing water and nutrients to the roots. Direct applications of the PTPNDS include plant space biology investigations on the Space Shuttle and plant research for life support in the Space Station Freedom. In the past, we investigated various configurations, the suitability of different porous materials, and the effects of pressure and pore size on plant growth. Current work is focused on characterizing the physical operation of the system, examining the effects of solution aeration, and developing prototype configurations for the Plant Growth Unit (PGU), the flight system for the Shuttle mid-deck. Future developments will involve testing on KC-135 parabolic flights, the design of flight hardware and testing aboard the Space Shuttle.

Investigation of Transonic Reynolds Number Scaling on a Twin-Engine Transport

NASA Technical Reports Server (NTRS)

Curtin, M. M.; Bogue, D. R.; Om, D.; Rivers, S. M. B.; Pendergraft, O. C., Jr.; Wahls, R. A.

2002-01-01

This paper discusses Reynolds number scaling for aerodynamic parameters including force and wing pressure measurements. A full-span model of the Boeing 777 configuration was tested at transonic conditions in the National Transonic Facility (NTF) at Reynolds numbers (based on mean aerodynamic chord) from 3.0 to 40.0 million. Data was obtained for a tail-off configuration both with and without wing vortex generators and flap support fairings. The effects of aeroelastics were separated from Reynolds number effects by varying total pressure and temperature independently. Data from the NTF at flight Reynolds number are compared with flight data to establish the wind tunnel/flight correlation. The importance of high Reynolds number testing and the need for developing a process for transonic Reynolds number scaling is discussed. This paper also identifies issues that need to be worked for Boeing Commercial to continue to conduct future high Reynolds number testing in the NTF.

Deployment and Evaluation of the Helicopter In-Flight Tracking System (HITS)

NASA Technical Reports Server (NTRS)

Daskalakis, Anastasios; Martone, Patrick

2004-01-01

The Gulf of Mexico airspace has two major operating regions: low altitude offshore (below 7,000 ft) and high altitude oceanic (above 18,000 ft). Both regions suffer significant inefficiencies due to the lack of continuous surveillance during Instrument Flight Rules operations. Provision of surveillance in the offshore region is hindered by its low-altitude nature, which makes coverage by conventional radars economically infeasible. Significant portions of the oceanic sectors are inaccessible to shore-based sensors, as they are beyond line-of-sight. Two emerging surveillance technologies were assessed that are relatively low cost and can be deployed on offshore platforms Wide Area Multilateration and Automatic Dependent Surveillance Broadcast. Performance criteria were formulated using existing FAA specifications. Three configurations were developed and deployed representative of systems serving full-size and reduced-sized domestic terminal areas and an en-route/oceanic region. These configurations were evaluated during nine flight test periods using fixed- and rotary-wing aircraft.

Lift cruise fan V/STOL aircraft conceptual design study T-39 modification. Volume 1: Technical report

NASA Technical Reports Server (NTRS)

Elliott, D. W.

1976-01-01

The conversion of two T-39 aircraft into lift cruise fan research and technology vehicles is discussed. The concept is based upon modifying the T-39A (NA265-40) Sabreliner airframe into a V/STOL configuration by incorporating two LCF-459 lift cruise fans and three YJ-97 gas generators. The propulsion concept provides the thrust for horizontal flight or lift for vertical flight by deflection of bifurcated nozzles while maintaining engine out safety throughout the flight envelope. The configuration meets all the study requirements specified for the design with control powers in VTOL and conversion in excess of the requirement making it an excellent vehicle for research and development. The study report consists of two volumes; Volume 1 (Reference a) contains background data detailed description and technical substantiation of the aircraft. Volume 2 includes cost data, scheduling and program planning not addressed in Volume 1.

Simulator study of flight characteristics of several large, dissimilar, cargo transport airplanes during approach and landing

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Smith, P. M.; Deal, P. L.; Neely, W. R., Jr.

1984-01-01

A six-degree-of-freedom, ground based simulator study is conducted to evaluate the low-speed flight characteristics of four dissimilar cargo transport airplanes. These characteristics are compared with those of a large, present-day (reference) transport configuration similar to the Lockheed C-5A airplane. The four very large transport concepts evaluated consist of single-fuselage, twin-fuselage, triple-fuselage, and span-loader configurations. The primary piloting task is the approach and landing operation. The results of his study indicate that all four concepts evaluated have unsatisfactory longitudinal and lateral directional low speed flight characteristics and that considerable stability and control augmentation would be required to improve these characteristics (handling qualities) to a satisfactory level. Through the use of rate command/attitude hold augmentation in the pitch and roll axes, and the use of several turn-coordination features, the handling qualities of all four large transports simulated are improved appreciably.

Wind-tunnel static and free-flight investigation of high-angle-of-attack stability and control characteristics of a model of the EA-6B airplane

NASA Technical Reports Server (NTRS)

Jordan, Frank L., Jr.; Hahne, David E.

1992-01-01

An investigation was conducted in the Langley 30- by 60-Foot Tunnel and the Langley 12-Foot Low-Speed Tunnel to identify factors contributing to a directional divergence at high angles of attack for the EA-6B airplane. The study consisted of static wind-tunnel tests, smoke and tuft flow-visualization tests, and free-flight tests of a 1/8.5-scale model of the airplane. The results of the investigation indicate that the directional divergence of the airplane is brought about by a loss of directional stability and effective dihedral at high angles of attack. Several modifications were tested that significantly alleviate the stability problem. The results of the free-flight study show that the modified configuration exhibits good dynamic stability characteristics and could be flown at angles of attack significantly higher than those of the unmodified configuration.

Bioinspired morphing wings for extended flight envelope and roll control of small drones

PubMed Central

Heitz, G.; Noca, F.; Floreano, D.

2017-01-01

Small-winged drones can face highly varied aerodynamic requirements, such as high manoeuvrability for flight among obstacles and high wind resistance for constant ground speed against strong headwinds that cannot all be optimally addressed by a single aerodynamic profile. Several bird species solve this problem by changing the shape of their wings to adapt to the different aerodynamic requirements. Here, we describe a novel morphing wing design composed of artificial feathers that can rapidly modify its geometry to fulfil different aerodynamic requirements. We show that a fully deployed configuration enhances manoeuvrability while a folded configuration offers low drag at high speeds and is beneficial in strong headwinds. We also show that asymmetric folding of the wings can be used for roll control of the drone. The aerodynamic performance of the morphing wing is characterized in simulations, in wind tunnel measurements and validated in outdoor flights with a small drone. PMID:28163882

Fielding An Amphibious UAV: Development, Results, and Lessons Learned

NASA Technical Reports Server (NTRS)

Pisanich, Greg; Morris, Stephen

2002-01-01

This report summarizes the work completed on the design and flight-testing of a small, unmanned, amphibious demonstrator aircraft that flies autonomously. The aircraft named ACAT (Autonomous Cargo Amphibious Transport) is intended to be a large cargo carrying unmanned aircraft that operates from water to avoid airspace and airfield conflict issues between manned and unmanned aircraft. To demonstrate the feasibility of this concept, a demonstrator ACAT was designed, built, and flown that has a six-foot wingspan and can fly autonomously from land or water airfield. The demonstrator was designed for a 1-hour duration and 1-mile telemetry range. A sizing code was used to design the smallest demonstrator UAV to achieve these goals. The final design was a six-foot wingspan, twin hull configuration that distributes the cargo weight across the span, reducing the wing structural weight. The demonstrator airframe was constructed from balsa wood, fiberglass, and plywood. A 4-stroke model airplane engine powered by methanol fuel was mounted in a pylon above the wing and powers the ACAT UAV. Initial flight tests from land and water were conducted under manual radio control and confirmed the amphibious capability of the design. Flight avionics that were developed by MLB for production UAVs were installed in the ACAT demonstrator. The flight software was also enhanced to permit autonomous takeoff and landing from water. A complete autonomous flight from ahard runway was successfully completed on July 5, 2001 and consisted of a take-off, rectangular flight pattern, and landing under complete computer control. A completely autonomous flight that featured a water takeoff and landing was completed on October 4, 2001. This report describes these activities in detail and highlights the challenges encountered and solved during the development of the ACAT demonstrator. hard runway was successfully completed on July 5, 2001 and consisted of a take-off, rectangular flight pattern, and landing under complete computer control. A completely autonomous flight that featured a water takeoff and landing was completed on October 4, 2001. This report describes these activities in detail and highlights the challenges encountered and solved during the development of the ACAT demonstrator.

High Stability Engine Control (HISTEC) Flight Test Results

NASA Technical Reports Server (NTRS)

Southwick, Robert D.; Gallops, George W.; Kerr, Laura J.; Kielb, Robert P.; Welsh, Mark G.; DeLaat, John C.; Orme, John S.

1998-01-01

The High Stability Engine Control (HISTEC) Program, managed and funded by the NASA Lewis Research Center, is a cooperative effort between NASA and Pratt & Whitney (P&W). The program objective is to develop and flight demonstrate an advanced high stability integrated engine control system that uses real-time, measurement-based estimation of inlet pressure distortion to enhance engine stability. Flight testing was performed using the NASA Advanced Controls Technologies for Integrated Vehicles (ACTIVE) F-15 aircraft at the NASA Dryden Flight Research Center. The flight test configuration, details of the research objectives, and the flight test matrix to achieve those objectives are presented. Flight test results are discussed that show the design approach can accurately estimate distortion and perform real-time control actions for engine accommodation.

Interface Supports Lightweight Subsystem Routing for Flight Applications

NASA Technical Reports Server (NTRS)

Lux, James P.; Block, Gary L.; Ahmad, Mohammad; Whitaker, William D.; Dillon, James W.

2010-01-01

A wireless avionics interface exploits the constrained nature of data networks in flight systems to use a lightweight routing method. This simplified routing means that a processor is not required, and the logic can be implemented as an intellectual property (IP) core in a field-programmable gate array (FPGA). The FPGA can be shared with the flight subsystem application. In addition, the router is aware of redundant subsystems, and can be configured to provide hot standby support as part of the interface. This simplifies implementation of flight applications requiring hot stand - by support. When a valid inbound packet is received from the network, the destination node address is inspected to determine whether the packet is to be processed by this node. Each node has routing tables for the next neighbor node to guide the packet to the destination node. If it is to be processed, the final packet destination is inspected to determine whether the packet is to be forwarded to another node, or routed locally. If the packet is local, it is sent to an Applications Data Interface (ADI), which is attached to a local flight application. Under this scheme, an interface can support many applications in a subsystem supporting a high level of subsystem integration. If the packet is to be forwarded to another node, it is sent to the outbound packet router. The outbound packet router receives packets from an ADI or a packet to be forwarded. It then uses a lookup table to determine the next destination for the packet. Upon detecting a remote subsystem failure, the routing table can be updated to autonomously bypass the failed subsystem.

The Orion Pad Abort 1 Flight Test A Highly Successful Test

NASA Technical Reports Server (NTRS)

Sinclair, Robert; Taylor, Anthony P. (Tony); Johnston, Justin

2011-01-01

The Orion Pad Abort 1 (PA-1) flight test was designed as an early demonstration of the Launch Abort System (LAS) for the Orion capsule. The LAS was designed developed and manufactured by the Lockheed Martin/Orbital Sciences team. At inception it was realized that recovery of the Orion Capsule simulator would be useful from an engineering analysis and data recovery point of view. Additionally this test represented a flight opportunity for the Orion parachute system, which in a real abort would provide final landing deceleration. The Orion parachute program is named CPAS (CEV Parachute Assembly System). Thus CPAS became a part of the PA-1 flight, as a secondary test objective. At program kick off, the CPAS system was in the design state described below. Airbag land landing of the spacecraft was the program baseline. This affected the rigging of the parachutes. The system entry deployment conditions and vehicle mass have both evolved since that original design. It was decided to use the baseline CPAS Generation 1 (Gen 1) parachute system for the recovery of the PA-1 flight. As CPAS was a secondary test objective, the system would be delivered in its developmental state. As the PA-1 program evolved, the parachute recovery system (CPAS) moved from a secondary objective to a more important portion of the program. Tests were added, weights and deployment conditions changed and some hardware portions of the CPAS configuration were not up to the new challenges. Additional tests were added to provide confidence in the developmental system. This paper will review a few of these aspects with the goal of showing some preliminary and qualitative results from what we believe was a highly successful test.

Flight-testing of the self-repairing flight control system using the F-15 highly integrated digital electronic control flight research facility

NASA Technical Reports Server (NTRS)

Stewart, James F.; Shuck, Thomas L.

1990-01-01

Flight tests conducted with the self-repairing flight control system (SRFCS) installed on the NASA F-15 highly integrated digital electronic control aircraft are described. The development leading to the current SRFCS configuration is highlighted. Key objectives of the program are outlined: (1) to flight-evaluate a control reconfiguration strategy with three types of control surface failure; (2) to evaluate a cockpit display that will inform the pilot of the maneuvering capacity of the damage aircraft; and (3) to flight-evaluate the onboard expert system maintenance diagnostics process using representative faults set to occur only under maneuvering conditions. Preliminary flight results addressing the operation of the overall system, as well as the individual technologies, are included.

Flight evaluation of pneumatic forebody vortex control in post-stall flight

NASA Technical Reports Server (NTRS)

Walchli, Lawrence A.

1994-01-01

The following topics are discussed: (1) X-29 description; Vortex Flow Control (VFC) technology description; (3) X-29 VFC wind tunnel results (forebody only); (4) X-29 VFC wind tunnel results (full configuration yawing moment); (5) X-29 VFC wind tunnel results (full configuration C(sub n) with sideslip); (6) X-29VFC wind tunnel results (full configuration pitching moment); (7) VFC optimized nozzle details; (8) X-29 forebody nozzle configuration; (9) X-29 VFC system stored gas schematic; (10) X-29 VFC system stored gas installation; (11) VFC effectiveness at zero sideslip; (12) VFC effectiveness at 35 AOA with sideslip; (13) 'VFC Roll' at 40 AOA; (14) Effects of VFC on wing rock; (15) Integrated controls C(sub n) prediction; (16) Proposed F-15 with lateral control laws with active VFC; (17) Simulated F-15 roll performance with active VFC; (18) Simulated F-15 spin recovery with active VFC; (19) Test team restructuring; (20) testbed selection; (21) Simulation for risk reduction; (22) Benefits of high pressure system; and (23) Advanced weapon system integration.

Titan/Centaur D-1T TC-2, Helios A flight data report. [of space missions

NASA Technical Reports Server (NTRS)

1975-01-01

Background data of spacecraft launching and flight are presented. A system analysis of the space vehicles is included, specifically on: (1) electronic equipment, (2) hydraulic equipment, (3) telemetry, (4) propulsion systems, (5) software (computers), and (6) guidance. Spacecraft and launch vehicle configurations are shown and described.

Time of flight mass spectrometer

DOEpatents

Ulbricht, Jr., William H.

1984-01-01

A time-of-flight mass spectrometer is described in which ions are desorbed from a sample by nuclear fission fragments, such that desorption occurs at the surface of the sample impinged upon by the fission fragments. This configuration allows for the sample to be of any thickness, and eliminates the need for complicated sample preparation.

ARC-1982-AC82-0253-11

NASA Image and Video Library

1982-04-07

Sikorsky Rotor Systems Research Aircraft ' RSRA' (72-001 NASA-740) compound configuration in flight: NASA Ames Research Center, Hangar and 40x 80x120ft W.T. in the background. Note: Used in publication in Flight Research at Ames; 57 Years of Development and Validation of Aeronautical Technology NASA SP-1998-3300 fig. 132

Aeroelastic Analysis for Rotorcraft in Flight or in a Wind Tunnel

NASA Technical Reports Server (NTRS)

Johnson, W.

1977-01-01

An analytical model is developed for the aeroelastic behavior of a rotorcraft in flight or in a wind tunnel. A unified development is presented for a wide class of rotors, helicopters, and operating conditions. The equations of motion for the rotor are derived using an integral Newtonian method, which gives considerable physical insight into the blade inertial and aerodynamic forces. The rotor model includes coupled flap-lag bending and blade torsion degrees of freedom, and is applicable to articulated, hingeless, gimballed, and teetering rotors with an arbitrary number of blades. The aerodynamic model is valid for both high and low inflow, and for axial and nonaxial flight. The rotor rotational speed dynamics, including engine inertia and damping, and the perturbation inflow dynamics are included. For a rotor on a wind-tunnel support, a normal mode representation of the test module, strut, and balance system is used. The aeroelastic analysis for the rotorcraft in flight is applicable to a general two-rotor aircraft, including single main-rotor and tandem helicopter configurations, and side-by-side or tilting proprotor aircraft configurations.

Boeing's variable geometry chevron: morphing aerospace structures for jet noise reduction

NASA Astrophysics Data System (ADS)

Calkins, Frederick T.; Mabe, James H.; Butler, George W.

2006-03-01

Boeing is applying cutting edge smart material actuators to the next generation morphing technologies for aircraft. This effort has led to the Variable Geometry Chevrons (VGC), which utilize compact, light weight, and robust shape memory alloy (SMA) actuators. These actuators morph the shape of chevrons on the trailing edge of a jet engine in order to optimize acoustic and performance objectives at multiple flight conditions. We have demonstrated a technical readiness level of 7 by successfully flight testing the VGCs on a Boeing 777-300ER with GE-115B engines. In this paper we describe the VGC design, development and performance during flight test. Autonomous operation of the VGCs, which did not require a control system or aircraft power, was demonstrated. A parametric study was conducted showing the influence of VGC configurations on shockcell generated cabin noise reduction during cruise. The VGC system provided a robust test vehicle to explore chevron configurations for community and shockcell noise reduction. Most importantly, the VGC concept demonstrated an exciting capability to optimize jet nozzle performance at multiple flight conditions.

A new flight control and management system architecture and configuration

NASA Astrophysics Data System (ADS)

Kong, Fan-e.; Chen, Zongji

2006-11-01

The advanced fighter should possess the performance such as super-sound cruising, stealth, agility, STOVL(Short Take-Off Vertical Landing),powerful communication and information processing. For this purpose, it is not enough only to improve the aerodynamic and propulsion system. More importantly, it is necessary to enhance the control system. A complete flight control system provides not only autopilot, auto-throttle and control augmentation, but also the given mission management. F-22 and JSF possess considerably outstanding flight control system on the basis of pave pillar and pave pace avionics architecture. But their control architecture is not enough integrated. The main purpose of this paper is to build a novel fighter control system architecture. The control system constructed on this architecture should be enough integrated, inexpensive, fault-tolerant, high safe, reliable and effective. And it will take charge of both the flight control and mission management. Starting from this purpose, this paper finishes the work as follows: First, based on the human nervous control, a three-leveled hierarchical control architecture is proposed. At the top of the architecture, decision level is in charge of decision-making works. In the middle, organization & coordination level will schedule resources, monitor the states of the fighter and switch the control modes etc. And the bottom is execution level which holds the concrete drive and measurement; then, according to their function and resources all the tasks involving flight control and mission management are sorted to individual level; at last, in order to validate the three-leveled architecture, a physical configuration is also showed. The configuration is distributed and applies some new advancement in information technology industry such line replaced module and cluster technology.

Research flight-control system development for the F-18 high alpha research vehicle

NASA Technical Reports Server (NTRS)

Pahle, Joseph W.; Powers, Bruce; Regenie, Victoria; Chacon, Vince; Degroote, Steve; Murnyak, Steven

1991-01-01

The F-18 high alpha research vehicle was recently modified by adding a thrust vectoring control system. A key element in the modification was the development of a research flight control system integrated with the basic F-18 flight control system. Discussed here are design requirements, system development, and research utility of the resulting configuration as an embedded system for flight research in the high angle of attack regime. Particular emphasis is given to control system modifications and control law features required for high angle of attack flight. Simulation results are used to illustrate some of the thrust vectoring control system capabilities and predicted maneuvering improvements.

Dryden Flight Research Center Overview

NASA Technical Reports Server (NTRS)

Meyer, Robert R., Jr.

2007-01-01

This viewgraph document presents a overview of the Dryden Flight Research Center's facilities. Dryden's mission is to advancing technology and science through flight. The mission elements are: perform flight research and technology integration to revolutionize aviation and pioneer aerospace technology, validate space exploration concepts, conduct airborne remote sensing and science observations, and support operations of the Space Shuttle and the ISS for NASA and the Nation. It reviews some of the recent research projects that Dryden has been involved in, such as autonomous aerial refueling, the"Quiet Spike" demonstration on supersonic F-15, intelligent flight controls, high angle of attack research on blended wing body configuration, and Orion launch abort tests.

Performance of a Supersonic Over-Wing Inlet with Application to a Low-Sonic-Boom Aircraft

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Hirt, Stefanie M.; Anderson, Bernhard H.; Fink, Lawrence E.; Magee, Todd E.

2014-01-01

Development of commercial supersonic aircraft has been hindered by many related factors including fuel-efficiency, economics, and sonic-boom signatures that have prevented over-land flight. Materials, propulsion, and flight control technologies have developed to the point where, if over-land flight were made possible, a commercial supersonic transport could be economically viable. Computational fluid dynamics, and modern optimization techniques enable designers to reduce the boom signature of candidate aircraft configurations to acceptable levels. However, propulsion systems must be carefully integrated with these low-boom configurations in order that the signatures remain acceptable. One technique to minimize the downward propagation of waves is to mount the propulsion systems above the wing, such that the wing provides shielding from shock waves generated by the inlet and nacelle. This topmounted approach introduces a number of issues with inlet design and performance especially with the highly-swept wing configurations common to low-boom designs. A 1.79%-scale aircraft model was built and tested at the NASA Glenn Research Center's 8-by 6-Foot Supersonic Wind Tunnel (8x6 SWT) to validate the configuration's sonic boom signature. In order to evaluate performance of the top-mounted inlets, the starboard flow-through nacelle on the aerodynamic model was replaced by a 2.3%-scale operational inlet model. This integrated configuration was tested at the 8x6 SWT from Mach 0.25 to 1.8 over a wide range of angles-of-attack and yaw. The inlet was also tested in an isolated configuration over a smaller range of angles-of-attack and yaw. A number of boundary-layer bleed configurations were investigated and found to provide a substantial positive impact on pressure recovery and distortion. Installed inlet performance in terms of mass capture, pressure recovery, and distortion over the Mach number range at the design angle-of-attack of 4-degrees is presented herein and compared to that at 0- degrees, as well as the isolated inlet configuration to highlight installation effects. Performance of the installed inlet fell below that of the isolated inlet at Mach numbers of 1.4 and greater. The installed inlet demonstrated adequate operability over the expected range of angles-of-attack and yaw, but did exhibit definite angle-ofattack and yaw limits at supersonic conditions. At each supersonic flight Mach number, performance parameters near zero yaw angle were relatively insensitive to yaw, but in general the yaw angle yielding best performance was non-zero and varied with angle-of-attack. Performance of the installed inlet is also presented as functions of angle-of-attack and yaw to highlight these effects. Distortion at the aerodynamic interface plane ranged between 10 and 25% at the inlet critical points over the range of flight Mach numbers tested and did not decrease significantly for the isolated inlet. Although these distortion levels would be considered high for operation with a turbine engine, the over-wing installation is likely not as significant a contributor as the low test Reynolds number. This is demonstrated by comparing CFD analysis of the isolated inlet at test scale with that at intermediate and full scales.

Space Shuttle 2 Advanced Space Transportation System. Volume 1: Executive Summary

NASA Technical Reports Server (NTRS)

Adinaro, James N.; Benefield, Philip A.; Johnson, Shelby D.; Knight, Lisa K.

1989-01-01

An investigation into the feasibility of establishing a second generation space transportation system is summarized. Incorporating successful systems from the Space Shuttle and technological advances made since its conception, the second generation shuttle was designed to be a lower-cost, reliable system which would guarantee access to space well into the next century. A fully reusable, all-liquid propellant booster/orbiter combination using parallel burn was selected as the base configuration. Vehicle characteristics were determined from NASA ground rules and optimization evaluations. The launch profile was constructed from particulars of the vehicle design and known orbital requirements. A stability and control analysis was performed for the landing phase of the orbiter's flight. Finally, a preliminary safety analysis was performed to indicate possible failure modes and consequences.

Calibration of the Verification Engineering Test Article-I (VETA-I) for AXAF using the VETA-I X-ray Detection System

NASA Technical Reports Server (NTRS)

Kellogg, E.; Brissenden, R.; Flanagan, K.; Freeman, M.; Hughes, J.; Jones, M.; Ljungberg, M.; Mckinnon, P.; Podgorski, W.; Schwartz, D.

1992-01-01

Advanced X-ray Astrophysics Facility (AXAF) X-ray optics testing is conducted by VETA-I, which consists of six nested Wolter type I grazing-incidence mirrors; VETA's X-ray Detection System (VXDS) in turn measures the imaging properties of VETA-I, yielding FWHM and encircled energy of the X-ray image obtained, as well as its effective area. VXDS contains a high resolution microchannel plate imaging X-ray detector and a pinhole scanning system in front of proportional-counter detectors. VETA-I's X-ray optics departs from the AXAF flight configuration in that it uses a temporary holding fixture; its mirror elements are not cut to final length, and are not coated with the metal film used to maximize high-energy reflection.

Direct solar heating for Space Station application

NASA Technical Reports Server (NTRS)

Simon, W. E.

1985-01-01

Early investigations have shown that a large percentage of the power generated on the Space Station will be needed in the form of high-temperature thermal energy. The most efficient method of satisfying this requirement is through direct utilization of available solar energy. A system concept for the direct use of solar energy on the Space Station, including its benefits to customers, technologists, and designers of the station, is described. After a brief discussion of energy requirements and some possible applications, results of selective tradeoff studies are discussed, showing area reduction benefits and some possible configurations for the practical use of direct solar heating. Following this is a description of system elements and required technologies. Finally, an assessment of available contributive technologies is presented, and a Space Shuttle Orbiter flight experiment is proposed.

Mobile 3d Mapping with a Low-Cost Uav System

NASA Astrophysics Data System (ADS)

Neitzel, F.; Klonowski, J.

2011-09-01

In this contribution it is shown how an UAV system can be built at low costs. The components of the system, the equipment as well as the control software are presented. Furthermore an implemented programme for photogrammetric flight planning and its execution are described. The main focus of this contribution is on the generation of 3D point clouds from digital imagery. For this web services and free software solutions are presented which automatically generate 3D point clouds from arbitrary image configurations. Possibilities of georeferencing are described whereas the achieved accuracy has been determined. The presented workflow is finally used for the acquisition of 3D geodata. On the example of a landfill survey it is shown that marketable products can be derived using a low-cost UAV.

Shuttle avionics and the goal language including the impact of error detection and redundancy management

NASA Technical Reports Server (NTRS)

Flanders, J. H.; Helmers, C. T.; Stanten, S. F.

1973-01-01

The relationship is examined between the space shuttle onboard avionics and the ground test computer language GOAL when used in the onboard computers. The study is aimed at providing system analysis support to the feasibility analysis of a GOAL to HAL translator, where HAL is the language used to program the onboard computers for flight. The subject is dealt with in three aspects. First, the system configuration at checkout, the general checkout and launch sequences, and the inventory of subsystems are described. Secondly, the hierarchic organization of onboard software and different ways of introducing GOAL-derived software onboard are described. Also the flow of commands and test data during checkout is diagrammed. Finally, possible impact of error detection and redundancy management on the GOAL language is discussed.

Eclipse - tow flight closeup and release

NASA Technical Reports Server (NTRS)

1998-01-01

This clip, running 15 seconds in length, shows the QF-106 'Delta Dart' gear down, with the tow rope secured to the attachment point above the aircraft nose. First there is a view looking back from the C-141A, then looking forward from the nose of the QF-106, and finally a shot of the aircraft being released from the tow rope. NASA Dryden Flight Research Center, Edwards, California, supported a Kelly Space and Technology, Inc. (KST)/U.S. Air Force project known as Eclipse, which demonstrated a reusable tow launch vehicle concept. The purpose of the project was to demonstrate a reusable tow launch vehicle concept that had been conceived and patented by KST. Kelly Space obtained a contract with the USAF Research Laboratory for the tow launch demonstration project under the Small Business Innovation Research (SBIR) program. The USAF SBIR contract included the modifications to turn the QF-106 into the Experimental Demonstrator #1 (EXD-01), and the C141A aircraft to incorporate the tow provisions to link the two aircraft, as well as conducting flight tests. The demonstration consisted of ground and flight tests. These tests included a Combined Systems Test of both airplanes joined by a tow rope, a towed taxi test, and six towed flights. The primary goal of the project was demonstrating the tow phase of the Eclipse concept using a scaled-down tow aircraft (C-141A) and a representative aerodynamically-shaped aircraft (QF-106A) as a launch vehicle. This was successfully accomplished. On December 20, 1997, NASA research pilot Mark Stucky flew a QF-106 on the first towed flight behind an Air Force C-141 in the joint Eclipse project with KST to demonstrate a reusable tow launch vehicle concept developed by KST. Kelly Space and Technology hoped to use the data from the tow tests to validate a tow-to-launch procedure for reusable space launch vehicles. Stucky flew six successful tow tests between December 1997 and February 6, 1998. On February 6, 1998, the sixth and final towed flight brought the project to a successful completion. Preliminary flight results determined that the handling qualities of the QF-106 on tow were very stable; actual flight-measured values of tow rope tension were well within predictions made by the simulation, aerodynamic characteristics and elastic properties of the tow rope were a significant component of the towing system; and the Dryden high-fidelity simulation provided a representative model of the performance of the QF-106 and C-141A airplanes in tow configuration. Total time on tow for the entire project was 5 hours, 34 minutes, and 29 seconds. All six flights were highly productive, and all project objectives were achieved. All three of the project objectives were successfully accomplished. The objectives were: demonstration of towed takeoff, climb-out, and separation of the EXD-01 from the towing aircraft; validation of simulation models of the towed aircraft systems; and development of ground and flight procedures for towing and launching a delta-winged airplane configuration safely behind a transport-type aircraft. NASA Dryden served as the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden also supplied engineering, simulation, instrumentation, range support, research pilots, and chase aircraft for the test series. Dryden personnel also performed the modifications to convert the QF-106 into the piloted EXD-01 aircraft. During the early flight phase of the project, Tracor, Inc. provided maintenance and ground support for the two QF-106 airplanes.The Air Force Flight Test Center (AFFTC), Edwards, California, provided the C-141A transport aircraft for the project, its flight and engineering support, and the aircrew. Kelly Space and Technology provided the modification design and fabrication of the hardware that was installed on the EXD-01 aircraft. Kelly Space and Technology hopes to use the data gleaned from the tow tests to develop a series of low-cost reusable launch vehicles, in particular to gain experience towing delta-wing aircraft having high wing loading, and in general to demonstrate various operational procedures such as ground processing and abort scenarios. The first successful towed flight occurred on Dec. 20, 1997. Prior to this first tow test flight, the C-141A and EXD-01 were used to conduct a series of tethered taxi tests to validate the tow procedures. Before these tethered taxi tests, a successful joint flight test was conducted in late October 1996, by Dryden, AFFTC, and KST, in which one of the Dryden F-18 chase aircraft flew at various ranges and locations behind the C-141A to define the wake turbulence and wingtip vortex environment. This flight test was replicated in July 1997, with an unmodified QF-106 flight proficiency aircraft.

V/STOL Systems Research Aircraft: A Tool for Cockpit Integration

NASA Technical Reports Server (NTRS)

Stortz, Michael W.; ODonoghue, Dennis P.; Tiffany, Geary (Technical Monitor)

1995-01-01

The next generation ASTOVL aircraft will have a complicated propulsion System. The configuration choices include Direct Lift, Lift-Fan and Lift+Lift /Cruise but the aircraft must also have supersonic performance and low-observable characteristics. The propulsion system may have features such as flow blockers, vectoring nozzles and flow transfer schemes. The flight control system will necessarily fully integrate the aerodynamic surfaces and the propulsive elements. With a fully integrated, fly-by-wire flight/propulsion control system, the options for cockpit integration are interesting and varied. It is possible to decouple longitudinal and vertical responses allowing the pilot to close the loop on flight path and flight path acceleration directly. In the hover, the pilot can control the translational rate directly without having to stabilize the inner rate and attitude loops. The benefit of this approach, reduced workload and increased precision. has previously been demonstrated through several motion-based simulations. In order to prove the results in flight, the V/STOL System Research Aircraft (VSRA) was developed at the NASA Ames Research Center. The VSRA is the YAV-8B Prototype modified with a research flight control system using a series-parallel servo configuration in all the longitudinal degrees of freedom (including thrust and thrust vector angle) to provide an integrated flight and propulsion control system in a limited envelope. Development of the system has been completed and flight evaluations of the response types have been performed. In this paper we will discuss the development of the VSRA, the evolution of the flight path command and translational rate command response types and the Guest Pilot evaluations of the system. Pilot evaluation results will be used to draw conclusions regarding the suitability of the system to satisfy V/STOL requirements.

Consensus of satellite cluster flight using an energy-matching optimal control method

NASA Astrophysics Data System (ADS)

Luo, Jianjun; Zhou, Liang; Zhang, Bo

2017-11-01

This paper presents an optimal control method for consensus of satellite cluster flight under a kind of energy matching condition. Firstly, the relation between energy matching and satellite periodically bounded relative motion is analyzed, and the satellite energy matching principle is applied to configure the initial conditions. Then, period-delayed errors are adopted as state variables to establish the period-delayed errors dynamics models of a single satellite and the cluster. Next a novel satellite cluster feedback control protocol with coupling gain is designed, so that the satellite cluster periodically bounded relative motion consensus problem (period-delayed errors state consensus problem) is transformed to the stability of a set of matrices with the same low dimension. Based on the consensus region theory in the research of multi-agent system consensus issues, the coupling gain can be obtained to satisfy the requirement of consensus region and decouple the satellite cluster information topology and the feedback control gain matrix, which can be determined by Linear quadratic regulator (LQR) optimal method. This method can realize the consensus of satellite cluster period-delayed errors, leading to the consistency of semi-major axes (SMA) and the energy-matching of satellite cluster. Then satellites can emerge the global coordinative cluster behavior. Finally the feasibility and effectiveness of the present energy-matching optimal consensus for satellite cluster flight is verified through numerical simulations.

STS-121 Mission Patch

NASA Image and Video Library

2005-06-01

STS121-S-001 (June 2005) --- The STS-121 patch depicts the space shuttle docked with the International Space Station (ISS) in the foreground, overlaying the astronaut symbol with three gold columns and a gold star. The ISS is shown in the configuration that it will be in during the STS-121 mission. The background shows the nighttime Earth with a dawn breaking over the horizon. STS-121, ISS mission ULF1.1, is the final Shuttle Return to Flight test mission. This utilization and logistics flight will bring a multipurpose logistics module (MPLM) to the ISS with several thousand pounds of new supplies and experiments. In addition, some new orbital replacement units (ORUs) will be delivered and stowed externally on ISS on a special pallet. These ORUs are spares for critical machinery located on the outside of the ISS. During this mission the crew will also carry out testing of shuttle inspection and repair hardware, as well as evaluate operational techniques and concepts for conducting on-orbit inspection and repair. The NASA insignia design for space shuttle flights is reserved for use by the astronauts and for other official use as the NASA Administrator may authorize. Public availability has been approved only in the forms of illustrations by the various news media. When and if there is any change in this policy, which is not anticipated, the change will be publicly announced. Photo credit: NASA

NASA's Space Launch System: SmallSat Deployment to Deep Space

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Creech, Stephen D.

2017-01-01

Leveraging the significant capability it offers for human exploration and flagship science missions, NASA's Space Launch System (SLS) also provides a unique opportunity for lower-cost deep-space science in the form of small-satellite secondary payloads. Current plans call for such opportunities to begin with the rocket's first flight; a launch of the vehicle's Block 1 configuration, capable of delivering 70 metric tons (t) to Low Earth Orbit (LEO), which will send the Orion crew vehicle around the moon and return it to Earth. On that flight, SLS will also deploy 13 CubeSat-class payloads to deep-space destinations. These secondary payloads will include not only NASA research, but also spacecraft from industry and international partners and academia. The payloads also represent a variety of disciplines including, but not limited to, studies of the moon, Earth, sun, and asteroids. While the SLS Program is making significant progress toward that first launch, preparations are already under way for the second, which will see the booster evolve to its more-capable Block 1B configuration, able to deliver 105t to LEO. That configuration will have the capability to carry large payloads co-manifested with the Orion spacecraft, or to utilize an 8.4-meter (m) fairing to carry payloads several times larger than are currently possible. The Block 1B vehicle will be the workhorse of the Proving Ground phase of NASA's deep-space exploration plans, developing and testing the systems and capabilities necessary for human missions into deep space and ultimately to Mars. Ultimately, the vehicle will evolve to its full Block 2 configuration, with a LEO capability of 130 metric tons. Both the Block 1B and Block 2 versions of the vehicle will be able to carry larger secondary payloads than the Block 1 configuration, creating even more opportunities for affordable scientific exploration of deep space. This paper will outline the progress being made toward flying smallsats on the first flight of SLS, and discuss future opportunities for smallsats on subsequent flights.

Technical Findings, Lessons Learned, and Recommendations Resulting from the Helios Prototype Vehicle Mishap

NASA Technical Reports Server (NTRS)

Noll, Thomas E.; Ishmael, Stephen D.; Henwood, Bart; Perez-Davis, Marla E.; Tiffany, Geary C.; Madura, John; Gaier, Matthew; Brown, John M.; Wierzbanowski, Ted

2007-01-01

The Helios Prototype was originally planned to be two separate vehicles, but because of resource limitations only one vehicle was developed to demonstrate two missions. The vehicle consisted of two configurations, one for each mission. One configuration, designated HP01, was designed to operate at extremely high altitudes using batteries and high-efficiency solar cells spread across the upper surface of its 247-foot wingspan. On August 13, 2001, the HP01 configuration reached an altitude of 96,863 feet, a world record for sustained horizontal flight by a winged aircraft. The other configuration, designated HP03, was designed for long-duration flight. The plan was to use the solar cells to power the vehicle's electric motors and subsystems during the day and to use a modified commercial hydrogen-air fuel cell system for use during the night. The aircraft design used wing dihedral, engine power, elevator control surfaces, and a stability augmentation and control system to provide aerodynamic stability and control. At about 30 minutes into the second flight of HP03, the aircraft encountered a disturbance in the way of turbulence and morphed into an unexpected, persistent, high dihedral configuration. As a result of the persistent high dihedral, the aircraft became unstable in a very divergent pitch mode in which the airspeed excursions from the nominal flight speed about doubled every cycle of the oscillation. The aircraft s design airspeed was subsequently exceeded and the resulting high dynamic pressures caused the wing leading edge secondary structure on the outer wing panels to fail and the solar cells and skin on the upper surface of the wing to rip away. As a result, the vehicle lost its ability to maintain lift, fell into the Pacific Ocean within the confines of the U.S. Navy's Pacific Missile Range Facility, and was destroyed. This paper describes the mishap and its causes, and presents the technical recommendations and lessons learned for improving the design, analysis, and testing methods and techniques required for this class of vehicle.

A prototype computerized synthesis methodology for generic space access vehicle (SAV) conceptual design

NASA Astrophysics Data System (ADS)

Huang, Xiao

2006-04-01

Today's and especially tomorrow's competitive launch vehicle design environment requires the development of a dedicated generic Space Access Vehicle (SAV) design methodology. A total of 115 industrial, research, and academic aircraft, helicopter, missile, and launch vehicle design synthesis methodologies have been evaluated. As the survey indicates, each synthesis methodology tends to focus on a specific flight vehicle configuration, thus precluding the key capability to systematically compare flight vehicle design alternatives. The aim of the research investigation is to provide decision-making bodies and the practicing engineer a design process and tool box for robust modeling and simulation of flight vehicles where the ultimate performance characteristics may hinge on numerical subtleties. This will enable the designer of a SAV for the first time to consistently compare different classes of SAV configurations on an impartial basis. This dissertation presents the development steps required towards a generic (configuration independent) hands-on flight vehicle conceptual design synthesis methodology. This process is developed such that it can be applied to any flight vehicle class if desired. In the present context, the methodology has been put into operation for the conceptual design of a tourist Space Access Vehicle. The case study illustrates elements of the design methodology & algorithm for the class of Horizontal Takeoff and Horizontal Landing (HTHL) SAVs. The HTHL SAV design application clearly outlines how the conceptual design process can be centrally organized, executed and documented with focus on design transparency, physical understanding and the capability to reproduce results. This approach offers the project lead and creative design team a management process and tool which iteratively refines the individual design logic chosen, leading to mature design methods and algorithms. As illustrated, the HTHL SAV hands-on design methodology offers growth potential in that the same methodology can be continually updated and extended to other SAV configuration concepts, such as the Vertical Takeoff and Vertical Landing (VTVL) SAV class. Having developed, validated and calibrated the methodology for HTHL designs in the 'hands-on' mode, the report provides an outlook how the methodology will be integrated into a prototype computerized design synthesis software AVDS-PrADOSAV in a follow-on step.

Integrated Application of Active Controls (IAAC) technology to an advanced subsonic transport project: Wing planform study and final configuration selection

NASA Technical Reports Server (NTRS)

1981-01-01

The Wing Planform Study and Final Configuration Selection Task of the Integrated Application of Active Controls (IAAC) Technology Project within the Energy Efficient Transport Program is documented. Application of Active Controls Technology (ACT) in combination with increased wing span resulted in significant improvements over the Conventional Baseline Configuration (Baseline) and the Initial ACT Configuration previously established. The configurations use the same levels of technology, takeoff gross weight, and payload as the Baseline. The Final ACT Configuration (Model 768-107) incorporates pitch-augmented stability (which enabled an approximately 10% aft shift in cruise center of gravity and a 44% reduction in horizontal tail size), lateral/directional-augmented stability, an angle-of-attack limiter, and wing-load alleviation. Flutter-mode control was not beneficial for this configuration. This resulted in an 890 kg (1960 lb) reduction in airplane takeoff gross weight and a 9.8% improvement in cruise lift/drag. At the Baseline mission range (3589 km 1938 nmi), this amounts to 10% block-fuel reduction. Results of this task strongly indicate that the IAAC Project should proceed with the Final ACT evaluation, and begin the required control system development and test.

Surface flow and heating distributions on a cylinder in near wake of Aeroassist Flight Experiment (AFE) configuration at incidence in Mach 10 Air

NASA Technical Reports Server (NTRS)

Wells, William L.

1990-01-01

Experimental heat transfer distributions and surface streamline directions are presented for a cylinder in the near wake of the Aeroassist Flight Experiment forebody configuration. Tests were conducted in air at a nominal free stream Mach number of 10, with post shock Reynolds numbers based on model base height of 6,450 to 50,770, and angles of attack of 5, 0, -5, and -10 degrees. Heat transfer data were obtained with thin film resistance gage and surface streamline directions by the oil flow technique. Comparisons between measured values and predicted values were made by using a Navier-Stokes computer code.

Microgravity Acceleration Measurement System (MAMS) Flight Configuration Verification and Status

NASA Technical Reports Server (NTRS)

Wagar, William

2000-01-01

The Microgravity Acceleration Measurement System (MAMS) is a precision spaceflight instrument designed to measure and characterize the microgravity environment existing in the US Lab Module of the International Space Station. Both vibratory and quasi-steady triaxial acceleration data are acquired and provided to an Ethernet data link. The MAMS Double Mid-Deck Locker (DMDL) EXPRESS Rack payload meets all the ISS IDD and ICD interface requirements as discussed in the paper which also presents flight configuration illustrations. The overall MAMS sensor and data acquisition performance and verification data are presented in addition to a discussion of the Command and Data Handling features implemented via the ISS, downlink and the GRC Telescience Center displays.

Users guide: The LaRC human-operator-simulator-based pilot model

NASA Technical Reports Server (NTRS)

Bogart, E. H.; Waller, M. C.

1985-01-01

A Human Operator Simulator (HOS) based pilot model has been developed for use at NASA LaRC for analysis of flight management problems. The model is currently configured to simulate piloted flight of an advanced transport airplane. The generic HOS operator and machine model was originally developed under U.S. Navy sponsorship by Analytics, Inc. and through a contract with LaRC was configured to represent a pilot flying a transport airplane. A version of the HOS program runs in batch mode on LaRC's (60-bit-word) central computer system. This document provides a guide for using the program and describes in some detail the assortment of files used during its operation.

Comparison ofdvanced turboprop interior noise control ground and flight test data

NASA Technical Reports Server (NTRS)

Simpson, Myles A.; Tran, Boi N.

1992-01-01

Interior noise ground tests conducted on a DC-9 aircraft test section are described. The objectives were to study ground test and analysis techniques for evaluating the effectiveness of interior noise control treatments for advanced turboprop aircraft, and to study the sensitivity of the ground test results to changes in various test conditions. Noise and vibration measurements were conducted under simulated advanced turboprop excitation, for two interior noise control treatment configurations. These ground measurement results were compared with results of earlier UHB (Ultra High Bypass) Demonstrator flight sts with comparable interior treatment configurations. The Demonstrator is an MD-80 test aircraft with the left JT8D engine replaced with a prototype UHB advanced turboprop engine.

Comparison ofdvanced turboprop interior noise control ground and flight test data

NASA Astrophysics Data System (ADS)

Simpson, Myles A.; Tran, Boi N.

Interior noise ground tests conducted on a DC-9 aircraft test section are described. The objectives were to study ground test and analysis techniques for evaluating the effectiveness of interior noise control treatments for advanced turboprop aircraft, and to study the sensitivity of the ground test results to changes in various test conditions. Noise and vibration measurements were conducted under simulated advanced turboprop excitation, for two interior noise control treatment configurations. These ground measurement results were compared with results of earlier UHB (Ultra High Bypass) Demonstrator flight sts with comparable interior treatment configurations. The Demonstrator is an MD-80 test aircraft with the left JT8D engine replaced with a prototype UHB advanced turboprop engine.

Simplified sonic-boom prediction. [using aerodynamic configuration charts and calculators or slide rules

NASA Technical Reports Server (NTRS)

Carlson, H. W.

1978-01-01

Sonic boom overpressures and signature duration may be predicted for the entire affected ground area for a wide variety of supersonic airplane configurations and spacecraft operating at altitudes up to 76 km in level flight or in moderate climbing or descending flight paths. The outlined procedure relies to a great extent on the use of charts to provide generation and propagation factors for use in relatively simple expressions for signature calculation. Computational requirements can be met by hand-held scientific calculators, or even by slide rules. A variety of correlations of predicted and measured sonic-boom data for airplanes and spacecraft serve to demonstrate the applicability of the simplified method.

The Use of Human Factors Simulation to Conserve Operations Expense

NASA Technical Reports Server (NTRS)

Hamilton, George S.; Dischinger, H. Charles, Jr.; Wu, Hsin-I.

1999-01-01

In preparation for on-orbit operations, NASA performs experiments aboard a KC-135 which performs parabolic maneuvers, resulting in short periods of microgravity. While considerably less expensive than space operations, the use of this aircraft is costly. Simulation of tasks to be performed during the flight can allow the participants to optimize hardware configuration and crew interaction prior to flight. This presentation will demonstrate the utility of such simulation. The experiment simulated is the fluid dynamics of epoxy components which may be used in a patch kit in the event of meteoroid damage to the International Space Station. Improved configuration and operational efficiencies were reflected in early and increased data collection.

Early Program Development

NASA Image and Video Library

1962-04-01

In this 1962 artist's concept , a proposed Nova rocket, shown at right, is compared to a Saturn C-1, left, and a Saturn C-5, center. The Marshall Space Flight Center directed studies of Nova configuration from 1960 to 1962 as a means of achieving a marned lunar landing with a direct flight to the Moon. Various configurations of the vehicle were examined, the largest being a five-stage vehicle using eight F-1 engines in the first stage. Although the program was effectively cancelled in 1962 when NASA planners selected the lunar-orbital rendezvous mode, the proposed F-1 engine was eventually used to propel the first stage of the Saturn V launch vehicle in the Apollo Program.

Comparison of several methods for estimating low speed stability derivatives

NASA Technical Reports Server (NTRS)

Fletcher, H. S.

1971-01-01

Methods presented in five different publications have been used to estimate the low-speed stability derivatives of two unpowered airplane configurations. One configuration had unswept lifting surfaces, the other configuration was the D-558-II swept-wing research airplane. The results of the computations were compared with each other, with existing wind-tunnel data, and with flight-test data for the D-558-II configuration to assess the relative merits of the methods for estimating derivatives. The results of the study indicated that, in general, for low subsonic speeds, no one text appeared consistently better for estimating all derivatives.

Trajectory optimization for A S.S.T.O. using in-flight LOX collection

NASA Astrophysics Data System (ADS)

Saint-Mard, M.; Hendrick, P.

A key point for a space mission (launch of a satellite, earth observation,…) is the optimization of the vehicle trajectory in order to burn the smallest quantity of propelant and then maximize the payload. This is true for evay space vehicle, but especially it is a crucial point for a Single-Stage-To-Orbit (SSTO) where the choice of a bad trajectory can result in an unrealizable vehicle due to the large airbreathing part of the flight In this study, we discuss the trajectory optimization for a Vertical Take-Off and Horizontal Landing (VTOHL) SSTO using supersonic in-flight atmospheric oxygen collection during a cruise phase (constant speed & constant altitude). This collected oxygen is stored in the LOX tanks and reused in the final rocket phase. This SSTO bas a Blended Body aerodynamic configuration as the one chosen by Lockheed Martin for its new space launcher (VentureStar and X-33). This SSTO uses rocket engines from take-off to Mach 1.7 and also for the exoatmospheric flight phase (that means for an altitude higher than 30km and a Mach number evolution from 6.8 to about 20). Between these two rocket phases, the SSTO is propelled by a subsonic ramjet. To perform this study, we use 2 computer programs (running on a home Computer): the first one allows to estimate the SSTO performances (TOGW, dry weight, hydrogen and oxygen consumptions) for a fixed payload mass and the second one permits the evaluation of the payload mass for a fixed TOGW.

Movable Ground Based Recovery System for Reuseable Space Flight Hardware

NASA Technical Reports Server (NTRS)

Sarver, George L. (Inventor)

2013-01-01

A reusable space flight launch system is configured to eliminate complex descent and landing systems from the space flight hardware and move them to maneuverable ground based systems. Precision landing of the reusable space flight hardware is enabled using a simple, light weight aerodynamic device on board the flight hardware such as a parachute, and one or more translating ground based vehicles such as a hovercraft that include active speed, orientation and directional control. The ground based vehicle maneuvers itself into position beneath the descending flight hardware, matching its speed and direction and captures the flight hardware. The ground based vehicle will contain propulsion, command and GN&C functionality as well as space flight hardware landing cushioning and retaining hardware. The ground based vehicle propulsion system enables longitudinal and transverse maneuverability independent of its physical heading.

Mission definition study for Stanford relativity satellite. Volume 2: Engineering flight test program

NASA Technical Reports Server (NTRS)

1971-01-01

The need is examined for orbital flight tests of gyroscope, dewar, and other components, in order to reduce the technical and financial risk in performing the relativity experiment. A program is described that would generate engineering data to permit prediction of final performance. Two flight tests are recommended. The first flight would test a dewar smaller than that required for the final flight, but of size and form sufficient to allow extrapolation to the final design. The second flight would use the same dewar design to carry a set of three gyroscopes, which would be evaluated for spinup and drift characteristics for a period of a month or more. A proportional gas control system using boiloff helium gas from the dewar, and having the ability to prevent sloshing of liquid helium, would also be tested.

Analysis and experimental studies of the control of hang gliders

NASA Technical Reports Server (NTRS)

Phillips, W. H.

1974-01-01

A theoretical analysis of the longitudinal and lateral characteristics of hang gliders in straight flight, pullups, and turns is presented. Some examples of the characteristics of a straight-wing configuration and a Rogallo-wing configuration are given. A means for improving the control of hang gliders while retaining the same basic control feel is proposed.

Flight set 360T004 (STS-30)

NASA Technical Reports Server (NTRS)

Nelsen, Lowell V.

1990-01-01

The performance of 360T004, Forth Flight, Redesigned Solid Rocket Motors (RSRM) is assessed in respect to joint sealing issues as seen from post-test inspection of the seals and sealing surfaces. The factory joint disassembly inspections for this flight set were omitted. The decision was based on the rational that there is sufficient information in the present data base, and this would give H-7 refurbishment operations faster turn around time for this set of hardware. The factory joint disassembly inspections will resume for 360H005, Fifth Flight, through 360L007, Seventh Flight, due to a new grease application being in effect during the assembly process. The left hand nozzle was forced into the snubbed position upon splash down. This required unique tooling to be manufactured to perform the disassembly of the internal nozzle joints. This was completed on February 5 and 6, 1990 at the H-5 Clearfield, Utah facility. The RSRM consisting of capture feature, field joints with the J-joint insulation configuration is illustrated. The nozzle-to-case joint design, which includes 100, 7/8-inch radial bolts in conjunction with a wiper O-ring and modified insulation design is also illustrated, as is the ignition system seals and a cross section of the igniter. The configuration of all internal nozzle joints is shown.

Computation of rotor aerodynamic loads in forward flight using a full-span free wake analysis

NASA Technical Reports Server (NTRS)

Quackenbush, Todd R.; Bliss, Donald B.; Wachspress, Daniel A.; Boschitsch, Alexander H.; Chua, Kiat

1990-01-01

The development of an advanced computational analysis of unsteady aerodynamic loads on isolated helicopter rotors in forward flight is described. The primary technical focus of the development was the implementation of a freely distorting filamentary wake model composed of curved vortex elements laid out along contours of constant vortex sheet strength in the wake. This model captures the wake generated by the full span of each rotor blade and makes possible a unified treatment of the shed and trailed vorticity in the wake. This wake model was coupled to a modal analysis of the rotor blade dynamics and a vortex lattice treatment of the aerodynamic loads to produce a comprehensive model for rotor performance and air loads in forward flight dubbed RotorCRAFT (Computation of Rotor Aerodynamics in Forward Flight). The technical background on the major components of this analysis are discussed and the correlation of predictions of performance, trim, and unsteady air loads with experimental data from several representative rotor configurations is examined. The primary conclusions of this study are that the RotorCRAFT analysis correlates well with measured loads on a variety of configurations and that application of the full span free wake model is required to capture several important features of the vibratory loading on rotor blades in forward flight.

Optical Component Performance for the Ocean Radiometer for Carbon Assessment (ORCA)

NASA Technical Reports Server (NTRS)

Quijada, Manuel A.; Wilson, Mark; Waluschka, Eugene; McClain, Charles R.

2011-01-01

The Ocean Radiometer for Carbon Assessment (ORCA) is a new design for the next generation remote sensing of ocean biology and biogeochemistry. ORCA is configured to meet all the measurement requirements of the Decadal Survey Aerosol, Cloud, and Ecology (ACE), the Ocean Ecosystem (OES) radiometer and the Pro-ACE climate data continuity mission (PACE). Under the auspices of a 2007 grant from NASA Research Opportunity in Space and Earth Science (ROSES) and the Instrument Incubator Program (IIP), a team at the Goddard Space Flight Center (GSFC) has been working on a functional prototype with flight-like fore and aft optics and scan mechanisms. As part of the development efforts to bring ORCA closer to a flight configuration and in order to reduce cost, we have conducted component-level optical testing using standard spectrophotometers and system-level characterizations using non-flight commercial off-the-shelf (COTS) focal plane array detectors. Although these arrays would not be able to handle flight data rates, they are adequate for optical alignment and performance testing. The purpose of this presentation is to describe the results of this testing performed at GSFC and the National Institute of Standards and Technology (NIST) at the component and system level. Specifically, we show results for ORCA's spectral calibration ranging from the near UV, visible, and near-infrared spectral region.

Functional integration of vertical flight path and speed control using energy principles

NASA Technical Reports Server (NTRS)

Lambregts, A. A.

1984-01-01

A generalized automatic flight control system was developed which integrates all longitudinal flight path and speed control functions previously provided by a pitch autopilot and autothrottle. In this design, a net thrust command is computed based on total energy demand arising from both flight path and speed targets. The elevator command is computed based on the energy distribution error between flight path and speed. The engine control is configured to produce the commanded net thrust. The design incorporates control strategies and hierarchy to deal systematically and effectively with all aircraft operational requirements, control nonlinearities, and performance limits. Consistent decoupled maneuver control is achieved for all modes and flight conditions without outer loop gain schedules, control law submodes, or control function duplication.

Ion beam plume and efflux characterization flight experiment study. [space shuttle payload

NASA Technical Reports Server (NTRS)

Sellen, J. M., Jr.; Zafran, S.; Cole, A.; Rosiak, G.; Komatsu, G. K.

1977-01-01

A flight experiment and flight experiment package for a shuttle-borne flight test of an 8-cm mercury ion thruster was designed to obtain charged particle and neutral particle material transport data that cannot be obtained in conventional ground based laboratory testing facilities. By the use of both ground and space testing of ion thrusters, the flight worthiness of these ion thrusters, for other spacecraft applications, may be demonstrated. The flight experiment definition for the ion thruster initially defined a broadly ranging series of flight experiments and flight test sensors. From this larger test series and sensor list, an initial flight test configuration was selected with measurements in charged particle material transport, condensible neutral material transport, thruster internal erosion, ion beam neutralization, and ion thrust beam/space plasma electrical equilibration. These measurement areas may all be examined for a seven day shuttle sortie mission and for available test time in the 50 - 100 hour period.

A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

NASA Technical Reports Server (NTRS)

Sim, Alex G.

1976-01-01

Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihooa estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5deg to 15.7deg. Data are presented for a subsonic and a transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

A Correlation Between Flight-Determined Derivatives and Wind-Tunnel Data for the X-24B Research Aircraft

NASA Technical Reports Server (NTRS)

Sim, Alex G.

1997-01-01

Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihood estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5 deg. to 15.7 deg. Data are presented for a subsonic and transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.

Potential Damage to Flight Hardware from MIL-STD-462 CS02 Setup

NASA Technical Reports Server (NTRS)

Harris, Patrick K.; Block, Nathan F.

2003-01-01

The MIL-STD-462 CS02 conducted susceptibility test setup includes an audio transformer, with the secondary used as an inductor, and a large capacitor. Together, these two components form an L-type low-pass filter to minimize the injected test signal input into the power source. Some flight hardware power input configurations are not compatible with this setup and break into oscillation when powered up. This, in turn, can damage flight hardware. Such an oscillation resulted in the catastrophic failure of an item tested in the Goddard Space Flight Center (GSFC) Large electromagnetic compatibility (EMC) Test Facility.

Potential Damage to Flight Hardware from MIL-STD-462 CS02 Setup

NASA Technical Reports Server (NTRS)

Harris, Patrick K.; Block, Nathan F.

2002-01-01

The MIL-STD-462 CS02 conducted susceptibility test setup, performed during electromagnetic compatibility (EMC) testing, consists of an audio transformer with the secondary used as an inductor and a large capacitor. Together, these two components form an L-type low-pass filter to minimize the injected test signal input into the power source. Some flight hardware power input configurations are not compatible with this setup and break into oscillation when powered up. This can damage flight hardware and caused a catastrophic failure to an item tested in the Goddard Space Flight Center (GSFC) Large EMC Test Facility.

Flight test techniques for validating simulated nuclear electromagnetic pulse aircraft responses

NASA Technical Reports Server (NTRS)

Winebarger, R. M.; Neely, W. R., Jr.

1984-01-01

An attempt has been made to determine the effects of nuclear EM pulses (NEMPs) on aircraft systems, using a highly instrumented NASA F-106B to document the simulated NEMP environment at the Kirtland Air Force Base's Vertically Polarized Dipole test facility. Several test positions were selected so that aircraft orientation relative to the test facility would be the same in flight as when on the stationary dielectric stand, in order to validate the dielectric stand's use in flight configuration simulations. Attention is given to the flight test portions of the documentation program.

DC-10-10 winglet flight test program management

NASA Technical Reports Server (NTRS)

Agar, J. R.

1982-01-01

This paper discusses the McDonnell Douglas/NASA DC-10-10 winglet flight test program from a program management viewpoint. The program was conducted to obtain flight test data on the same airplane with and without winglets for direct comparison. As occasionally happens in flight tests, unexpected events occur. This program was encumbered by a low-speed buffet anomaly that required several configuration modifications before satisfactory performance could be attained. This paper relates the management techniques utilized to accommodate the unplanned increases in program scope and still complete the program on time and below the budgeted cost.

Identification of pilot dynamics from in-flight tracking data

NASA Technical Reports Server (NTRS)

Hess, R. A.; Mnich, M. A.

1985-01-01

Data from a representative flight task involving an F-14 'pursuer' aircraft tracking a T-38 'target' aircraft in a 3G wind-up turn and in level flight are processed using a least squares identification technique in an attempt to identify pilot/vehicle dynamics. Comparative identification results are provided by a Fourier coefficient method which requires a carefully designed and implemented input consisting of a sum of sinusoids. The least-squares results compare favorably with those obtained by the Fourier technique. An example of crossover frequency regression is discussed in the light of the conditions of one of the flight configurations.

ASCAN Helms inspects T-38A during flight training at Ellington Field

NASA Technical Reports Server (NTRS)

1990-01-01

1990 Group 13 Astronaut Candidate (ASCAN) Susan J. Helms inspects a T-38A cockpit configuration during flight training at Ellington Field. Helms, along with other 1990 Astronaut Class members, participated in the training conducted by NASA staff pilots on 07-26-90 and 07-27-90. Ellington Field is located near JSC.

Booster Separation Motor (BSM) Test Fire

NASA Technical Reports Server (NTRS)

2007-01-01

This photograph depicts a hot fire test of the Shuttle Booster Separation Motor (BSM) at the Marshall Space Flight Center (MSFC) test stand 116. The objective of the test was to test the aft heat seal in flight configuration. The function of the motor is to separate the Shuttle vehicle from the boosters that carry it into space.

Results of a simulator test comparing two display concepts for piloted flight-path-angle control

NASA Technical Reports Server (NTRS)

Kelley, W. W.

1978-01-01

Results of a simulator experiment which was conducted in order to compare pilot gamma-control performance using two display formats are reported. Pilots flew a variable flight path angle tracking task in the landing configuration. Pilot and airplane performance parameters were recorded and pilot comments noted for each case.

Room-temperature quantum noise limited spectrometry and methods of the same

DOE Office of Scientific and Technical Information (OSTI.GOV)

Stevens, Charles G.; Tringe, Joseph W.; Cunningham, Christopher T.

According to one embodiment, a heterodyne detection system for detecting light, includes: a first input aperture configured to receive first light from a scene input; a second input aperture configured to receive second light from a local oscillator input; a broadband local oscillator configured to provide the second light to the second input aperture; a dispersive element configured to disperse the first light and the second light; and a final condensing lens coupled to a detector. The final condensing lens is configured to concentrate incident light from a primary condensing lens onto the detector. The detector is configured to sensemore » a frequency difference between the first light and the second light; and the final condensing lens comprises a plasmonic condensing lens. Methods for forming a plasmonic condensing lens to enable room temperature quantum noise limited spectrometry are also disclosed.« less

Final Environmental Assessment for Low-Level Flight Testing, Evaluation, and Training, Edwards Air Force Base

DTIC Science & Technology

2005-05-01

4. TITLE AND SUBTITLE Final Environmental Assessment for Low-Level Flight Testing, Evaluation, and Training, Edwards Air Force Base 5a. CONTRACT...NAME(S) AND ADDRESS(ES) Air Force Flight Test Center,Environmental Management Directorate,Edwards AFB,CA,93524 8. PERFORMING ORGANIZATION REPORT...DISTRIBUTION/AVAILABILITY STATEMENT Approved for public release; distribution unlimited 13. SUPPLEMENTARY NOTES 14. ABSTRACT The U.S. Air Force Flight Test

AFTI/F-16 50th flight team photo

NASA Technical Reports Server (NTRS)

1983-01-01

An early (1983) photograph of the AFTI F-16 team, commemorating the aircraft's 50th flight. It shows the initial configuration and paint finish of the AFTI F-16, as well as the forward mounted canards and the spin chute. During the 1980s and 1990s, NASA and the U.S. Air Force participated in a joint program to integrate and demonstrate new avionics technologies to improve close air support capabilities in next-generation aircraft. The testbed aircraft, seen here in flight over the desert at NASA's Dryden Flight Research Center, Edwards, California, was called the Advanced Fighter Technology Integration (AFTI) F-16. The tests demonstrated technologies to improve navigation and the pilot's ability to find and destroy enemy ground targets day or night, including adverse weather. The aircraft--an F-16A Fighting Falcon (Serial #75-0750)--underwent numerous modifications. A relatively low-cost testbed, it evaluated the feasability of advanced, intergrated-sensor, avionics, and flight control technologies. During the first phase of the AFTI/F-16 program, which began in 1983, the aircraft demonstrated voice-actuated commands, helmet-mounted sights, flat turns, and selective fuselage pointing using forward-mounted canards and a triplex digital flight control computer system. The second phase of research, which began in the summer of 1991, demonstrated advanced technologies and capabilities to find and destroy ground targets day or night, and in adverse weather while using maneuverability and speed at low altitude. This phase was known as the close air support and battlefield air interdiction (CAS/BAI) phase. Finally, the aircraft was used to assess the Automatic Ground Collision Avoidance System (Auto - GCAS), a joint project with the Swedish Government. For these tests, the pilot flew the aircraft directly toward the ground, simulating a total loss of control. The GCAS was designed to take command in such emergencies and bring the aircraft back to level flight. The AFTI F-16 program ended at Dryden on November 4, 1997 after 15 years and over 700 research flights. The USAF continued to fly the aircraft until retiring it to the Air Force Museum on January 9, 2001.

Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles, volume 1. [supersonic cruise aircraft research wind tunnel tests

NASA Technical Reports Server (NTRS)

Kozlowski, H.; Packman, A. B.

1978-01-01

Jet noise spectra obtained at static conditions from an acoustic wind tunnel and an outdoor facility are compared. Data curves are presented for (1) the effect of relative velocity on OASPL directivity (all configurations); (2) the effect of relative velocity on noise spectra (all configurations); (3) the effect of velocity on PNL directivity (coannular nozzle configurations); (4) nozzle exhaust plume velocity profiles; and (5) the effect of relative velocity on aerodynamic performance.