Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “test data”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,009 records · Page 56
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Viking entry aerodynamics and heating

The characteristics of the Mars entry including the mission sequence of events and associated spacecraft weights are described along with the Viking spacecraft. Test data are presented for the aerodynamic characteristics of the entry vehicle showing trimmed alpha, drag coefficient, and trimmed lift to drag ratio versus Mach number; the damping characteristics of the entry configuration; the angle of attack time history of Viking entries; stagnation heating and pressure time histories; and the aeroshell heating distribution as obtained in tests run in a shock tunnel for various gases. Flight tests which demonstrate the aerodynamic separation of the full-scale aeroshell and the flying qualities of the entry configuration in an uncontrolled mode are documented. Design values selected for the heat protection system based on the test data and analysis performed are presented.

Polutchko, R. J.↗
NASA NTRS2005

Empirical Prediction of Aircraft Landing Gear Noise

This report documents a semi-empirical/semi-analytical method for landing gear noise prediction. The method is based on scaling laws of the theory of aerodynamic noise generation and correlation of these scaling laws with current available test data. The former gives the method a sound theoretical foundation and the latter quantitatively determines the relations between the parameters of the landing gear assembly and the far field noise, enabling practical predictions of aircraft landing gear noise, both for parametric trends and for absolute noise levels. The prediction model is validated by wind tunnel test data for an isolated Boeing 737 landing gear and by flight data for the Boeing 777 airplane. In both cases, the predictions agree well with data, both in parametric trends and in absolute noise levels.

Golub, Robert A.↗
NASA NTRS1989

NASA Lewis Stirling engine computer code evaluation

In support of the U.S. Department of Energy's Stirling Engine Highway Vehicle Systems program, the NASA Lewis Stirling engine performance code was evaluated by comparing code predictions without engine-specific calibration factors to GPU-3, P-40, and RE-1000 Stirling engine test data. The error in predicting power output was -11 percent for the P-40 and 12 percent for the Re-1000 at design conditions and 16 percent for the GPU-3 at near-design conditions (2000 rpm engine speed versus 3000 rpm at design). The efficiency and heat input predictions showed better agreement with engine test data than did the power predictions. Concerning all data points, the error in predicting the GPU-3 brake power was significantly larger than for the other engines and was mainly a result of inaccuracy in predicting the pressure phase angle. Analysis into this pressure phase angle prediction error suggested that improvements to the cylinder hysteresis loss model could have a significant effect on overall Stirling engine performance predictions.

Sullivan, Timothy J.↗
NASA NTRS1982

Pressure distribution data from tests of 2.29-meter (7.5-ft.) span EET high-lift research model in Langley 4- by 7-meter tunnel

A 2.29 m (7.5 ft.) span high-lift research model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Langley 4- by 7-Meter Tunnel to determine the low speed performance characteristics of a representative high aspect ratio suprcritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.

Morgan, H. L., Jr.↗
NASA NTRS1983

Pressure distribution data from tests of 2.29 M (7.5 feet) span EET high-lift transport aircraft model in the Ames 12-foot pressure tunnel

A high-lift transport aircraft model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Ames 12-ft pressure tunnel to determine the low-speed performance characteristics of a representative high-aspect-ratio supercritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.

Kjelgaard, S. O.↗
NASA NTRS2002

Use of TRMM Data to Test an Improved Parameterization of Stratiform Precipitation

We successfully implemented our EAUcup parameterization of convection in our most recent geodesic-grid version of the Colorado State University General Circulation Model (CSU GCM). Tendencies due to convection and large-scale cloud microphysics are computed using the same time-step, allowing smoother interactions between convective and large-scale moist processes. Figures illustrate our most recent progress on the simulation of convective and stratiform precipitations for January. Comparisons against TRMM precipitation radar (PR) data would reveal a systematic overestimation of the simulated monthly-mean convective rainfall over well-known areas of deep convection. One major deficiency in the global distribution of stratiform precipitation simulated in the CSU GCM is the lack of stratiform precipitation over land in the tropics. This result strongly disagree with Fowler and Randall who show that convective and stratiform precipitations always coexist and that their ratio to the total precipitation is about 0.5. We believe that stratiform precipitation over land evaporates before reaching the ground because the lower troposphere remains too dry in response to the lack of convective detrainment occurring immediately above the Planetary Boundary Layer (PBL). This issue is further investigated at present. We made significant progress in the development of our parameterization of fractional cloudiness. As an example of our ongoing research effort, a figure shows the global distribution of the cloud ice mixing ratio and corresponding fractional area of ice clouds at 250 hPa. These results were obtained using our latitude-longitude grid version of the CSU GCM. There is a nice correspondence between areas of high (low) values of the cloud ice mixing ratio and the cloud fraction. A nice correspondence between the global distribution of the cloud water mixing ratio and fraction of water clouds in the lower troposphere can also be found. We are pursuing this effort, but in the geodesic version of the GCM. We hope to be able to make comparison of the simulated precipitation against TRMM precipitation data in the near future.

Randall, David A.↗
NASA NTRS1980

Convective heating of solid propellant during ignition

A pattern search optimization procedure is used to deduce from test data on the solid rocket boosters of the Space Shuttle the convective heat transfer coefficients between the igniter and main motor combustion gases and the surface of the solid propellant by utilizing a spatial and temporal ignition transient analysis. The 'best' convective coefficients are found which will minimize the overall deviation of the analytical prediction from the ignition transient test data. Computationally, this is achieved by treating the motor in such a way that only four optimization variables must be determined. The resulting coefficients greatly improve the prediction capabilities of the ignition transient analysis for the Space Shuttle SRM.

Murph, J. E.↗
NASA NTRS2000

Unsteady Aerodynamic Effects on the Flight Characteristics of an F-16XL Configuration

Unsteady aerodynamic models based on windtunnel forced oscillation test data and analyzed with a fuzzy logic algorithm arc incorporated into an F-16XL flight simulation code. The reduced frequency needed in the unsteady models is numerically calculated by using a limited prior time history of state variables in a least-square sense. Numerical examples arc presented to show the accuracy of the calculated reduced frequency. Oscillatory control inputs are employed to demonstrate the differences in the flight characteristics based on unsteady and quasi-steady aerodynamic models. Application of the unsteady aerodynamic models is also presented and the results are compared with one set of F16XIL longitudinal maneuver flight data. It is shown that the main differences in dynamic response are in the lateral-directional characteristics, with the quasi-steady model being more stable than the flight vehicle, while the unsteady model being more unstable. Similar conclusions can also be made in a simulated rapid sideslipping roll. To improve unsteady aerodynamic modeling, it is recommended to acquire test data with coupled motions in pitch, roll and yaw.

Wang, Zhongjun↗
NASA NTRS1997

Improved Hypersonic Inlet Performance Using Validated Strut Compression Designs

Aerojet is currently executing two Strutjet propulsion contracts: one a Rocket Based Combined Cycle (RBCC) engine for a NASA-Marshall Space Flight Center (MSFC) Advanced Reusable Transportation Technology (ARTT) program, the second a Dual Mode Ram/Scramjet engine for a USAF Wright Laboratories Storable Fuel Scramjet Flow Path Concepts program. The engines employed in both programs operate at supersonic and low hypersonic speeds and use inlets employing forebody external and sidewall compression. Aerojet has developed and validated a successful design methodology applicable to these inlet types. Design features include an integrated vehicle forebody, external side compression struts, strut sidewall and throat bleed, a throat shock trap, and variable geometry internal contraction. Computation Fluid Dynamic (CFD) predictions and test data show these inlets allow substantially increased flow turning angles over other designs. These increased flow turning angles allow shorter and lighter engines than current designs, which in turn enables higher performing vehicles with broad operating characteristics. This paper describes the designs of two different inlets evaluated by the NASA-MSFC and USAF programs, discusses the results of wind tunnel tests performed by NASA-Lewis Research Center, and provides correlations of test data with CFD predictions. Parameters of interest include low Mach number starting capability, start sensitivity as a function of back pressure at various contraction ratios, flow turning angles, strut and throat bleed effects, and pressure recovery at various Mach numbers.

Bulman, M. J.↗
NASA NTRS2005

Seeing the Invisible: Embedding Tests in Code That Cannot be Modified

The difficulty of characterizing and observing valid software behavior during testing can be very difficult in flight systems. To address this issue, we evaluated several approaches to increasing test observability on the Shuttle Abort Flight Management (SAFM) system. To increase test observability, we added probes into the running system to evaluate the internal state and analyze test data. To minimize the impact of the instrumentation and reduce manual effort, we used Aspect-Oriented Programming (AOP) tools to instrument the source code. We developed and elicited a spectrum of properties, from generic to application specific properties, to be monitored via the instrumentation. To evaluate additional approaches, SAFM was ported to Linux, enabling the use of gcov for measuring test coverage, Valgrind for looking for memory usage errors, and libraries for finding non-normal floating point values. An in-house C++ source code scanning tool was also used to identify violations of SAFM coding standards, and other potentially problematic C++ constructs. Using these approaches with the existing test data sets, we were able to verify several important properties, confirm several problems and identify some previously unidentified issues.

O'Malley, Owen↗
NASA NTRS1999

System Identification for Integrated Aircraft Development and Flight Testing [l'Identification Des Systemes Pour le Developpement Integre des Aeronefs et les Essais en Vol]

Over the last decades flight vehicles such as aircraft and helicopters entering service and requiring increased operational effectiveness have with few exceptions experienced prolonged flight test development to achieve full certification. In many cases the original requirements had later to be reduced to enable release to service. The impact on the customer, and manufacturer has been considerable leading to increased costs and or reduced operational capabilities. These costly experiences are largely a result of the flight vehicle not behaving as modelled and designed. The evaluation of flight test data can be used as a tool for validating windtunnel results and mathematical models describing the flight dynamical behaviour. In this sense the uncertainty of important aerodynamic stability and control parameters can be reduced and the confidence of aircraft mathematical models improved. An additional important factor comes from the implementation of active control systems offering the promise of significantly increased flight vehicle performance and operational capability. This approach extends the traditional trade-offs between aerodynamics, structures and propulsion systems to include full- time, full-authority fly-by-wire/light systems. It is imperative that the aerodynamic stability and control parameters of such integrated flight and propulsion control systems have to turn out inflight as predicted, since inherent stability margins will be lower and the flight control system must correct these deficiencies to provide flight critical redundancy and safety. With the methodology of system identification from flight tests it is possible to sense the control inputs and the flight vehicle reactions Such as accelerations, rates and attitudes. The mathematical model, e.g. the model structure and parameters, has to be determined from the relationship of the measured control inputs and the system's responses. The aim of this symposium was to review the present state of the art of flight vehicle system and parameter identification techniques, and to provide a critical appraisal of current methods developed and applied to flight test data in a number of NATO nations. Particular emphasis was placed on practical aspects and lessons learned in order to generate information useful to the flight test community in industry and government agencies. The technical papers share invaluable experience and emphasize the advances of flight vehicle system identification over the last years to the point where confidence and robustness level is now reasonably high. The symposium covered overviews of identification methodologies, flight test techniques, recent aircraft and helicopter application programs, and a session of short papers covering up-to-the-minute flight test results. A final discussion included prepared comments from experts and concluded with key issues learned in the application of system identification and future research needs. The essential benefits to NATO nations can be condensed as follows: More accurate mathematical models for high bandwidth flight control systems, Improved assessment and evaluation of flying qualities, High fidelity mathematical models for flight vehicle development and mission training simulators, and generally, Reduced flight test time and costs.

Advisory Group for Aerospace Research and Developm↗
NASA NTRS1975

A collection of articles on S/X-band experiment zero delay ranging tests, volume 1

Articles are presented which are concerned with the development of special test equipment and a dual-frequency zero delay device (ZDD) that were required for range tests and the measurement of ground station delays for the Mariner-Venus-Mercury 1973 S/X-band experiment. Test data obtained at DSS 14 after installation of the ZDD on the 64-m antenna are given. It is shown that large variations of range were observed as a function of antenna elevation angle and were sensitive to antenna location. A ranging calibration configuration that was subsequently developed and a technique for determining the appropriate Z-correction are described. Zero delay test data at DSS 14 during the Mariner 10 Venus-Mercury-Encounter periods (1974 days 12-150) are presented. The theoretical analysis and experimental verifications are included of the effects of multipath and effects of discontinuities on range delay measurements. A movable subreflector technique and the multipath theory were used to isolate principal multipath errors on the 64-m antenna and to enable a more accurate determination of the actual ground station range delay.

Otoshi, T. Y.↗
NASA NTRS2016

Thermal Vacuum Test Correlation of a Zero Propellant Load Case Thermal Capacitance Propellant Gauging Analytical Model

This thesis describes the development and correlation of a thermal model that forms the foundation of a thermal capacitance spacecraft propellant load estimator. Specific details of creating the thermal model for the diaphragm propellant tank used on NASA's Magnetospheric Multiscale spacecraft using ANSYS and the correlation process implemented are presented. The thermal model was correlated to within plus or minus 3 degrees Celsius of the thermal vacuum test data, and was determined sufficient to make future propellant predictions on MMS. The model was also found to be relatively sensitive to uncertainties in applied heat flux and mass knowledge of the tank. More work is needed to improve temperature predictions in the upper hemisphere of the propellant tank where predictions were found to be 2 to 2.5 C lower than the test data. A road map for applying the model to predict propellant loads on the actual MMS spacecraft toward its end of life in 2017-2018 is also presented.

Gauging↗
NASA NTRS1993

An analytic modeling and system identification study of rotor/fuselage dynamics at hover

A combination of analytic modeling and system identification methods have been used to develop an improved dynamic model describing the response of articulated rotor helicopters to control inputs. A high-order linearized model of coupled rotor/body dynamics including flap and lag degrees of freedom and inflow dynamics with literal coefficients is compared to flight test data from single rotor helicopters in the near hover trim condition. The identification problem was formulated using the maximum likelihood function in the time domain. The dynamic model with literal coefficients was used to generate the model states, and the model was parametrized in terms of physical constants of the aircraft rather than the stability derivatives resulting in a significant reduction in the number of quantities to be identified. The likelihood function was optimized using the genetic algorithm approach. This method proved highly effective in producing an estimated model from flight test data which included coupled fuselage/rotor dynamics. Using this approach it has been shown that blade flexibility is a significant contributing factor to the discrepancies between theory and experiment shown in previous studies. Addition of flexible modes, properly incorporating the constraint due to the lag dampers, results in excellent agreement between flight test and theory, especially in the high frequency range.

Hong, Steven W.↗
NASA NTRS1993

An analytic modeling and system identification study of rotor/fuselage dynamics at hover

A combination of analytic modeling and system identification methods have been used to develop an improved dynamic model describing the response of articulated rotor helicopters to control inputs. A high-order linearized model of coupled rotor/body dynamics including flap and lag degrees of freedom and inflow dynamics with literal coefficients is compared to flight test data from single rotor helicopters in the near hover trim condition. The identification problem was formulated using the maximum likelihood function in the time domain. The dynamic model with literal coefficients was used to generate the model states, and the model was parametrized in terms of physical constants of the aircraft rather than the stability derivatives, resulting in a significant reduction in the number of quantities to be identified. The likelihood function was optimized using the genetic algorithm approach. This method proved highly effective in producing an estimated model from flight test data which included coupled fuselage/rotor dynamics. Using this approach it has been shown that blade flexibility is a significant contributing factor to the discrepancies between theory and experiment shown in previous studies. Addition of flexible modes, properly incorporating the constraint due to the lag dampers, results in excellent agreement between flight test and theory, especially in the high frequency range.

Hong, Steven W.↗
NASA NTRS1978

Space Shuttle Plume Simulation Effect on Aerodynamics

Technology for simulating plumes in wind tunnel tests was not adequate to provide the required confidence in test data where plume induced aerodynamic effects might be significant. A broad research program was undertaken to correct the deficiency. Four tasks within the program are reported. Three of these tasks involve conducting experiments, related to three different aspects of the plume simulation problem: (1) base pressures; (2) lateral jet pressures; and (3) plume parameters. The fourth task involves collecting all of the base pressure test data generated during the program. Base pressures were measured on a classic cone ogive cylinder body as affected by the coaxial, high temperature exhaust plumes of a variety of solid propellant rockets. Valid data were obtained at supersonic freestream conditions but not at transonic. Pressure data related to lateral (separation) jets at M infinity = 4.5, for multiple clustered nozzles canted to the freestream and operating at high dynamic pressure ratios. All program goals were met although the model hardware was found to be large relative to the wind tunnel size so that operation was limited for some nozzle configurations.

Hair, L. M.↗
NASA NTRS2010

Reliability Quantification of Advanced Stirling Convertor (ASC) Components

The Advanced Stirling Convertor, is intended to provide power for an unmanned planetary spacecraft and has an operational life requirement of 17 years. Over this 17 year mission, the ASC must provide power with desired performance and efficiency and require no corrective maintenance. Reliability demonstration testing for the ASC was found to be very limited due to schedule and resource constraints. Reliability demonstration must involve the application of analysis, system and component level testing, and simulation models, taken collectively. Therefore, computer simulation with limited test data verification is a viable approach to assess the reliability of ASC components. This approach is based on physics-of-failure mechanisms and involves the relationship among the design variables based on physics, mechanics, material behavior models, interaction of different components and their respective disciplines such as structures, materials, fluid, thermal, mechanical, electrical, etc. In addition, these models are based on the available test data, which can be updated, and analysis refined as more data and information becomes available. The failure mechanisms and causes of failure are included in the analysis, especially in light of the new information, in order to develop guidelines to improve design reliability and better operating controls to reduce the probability of failure. Quantified reliability assessment based on fundamental physical behavior of components and their relationship with other components has demonstrated itself to be a superior technique to conventional reliability approaches based on utilizing failure rates derived from similar equipment or simply expert judgment.

Shah, Ashwin R.↗
NASA NTRS2010

Acoustic-Modal Testing of the Ares I Launch Abort System Attitude Control Motor Valve

The Attitude Control Motor (ACM) is being developed for use in the Launch Abort System (LAS) of NASA's Ares I launch vehicle. The ACM consists of a small solid rocket motor and eight actuated pintle valves that directionally allocate.thrust_- 1t.has-been- predicted-that significant unsteady. pressure.fluctuations.will.exist. inside the-valves during operation. The dominant frequencies of these oscillations correspond to the lowest several acoustic natural frequencies of the individual valves. An acoustic finite element model of the fluid volume inside the valve has been critical to the prediction of these frequencies and their associated mode shapes. This work describes an effort to experimentally validate the acoustic finite model of the valve with an acoustic modal test. The modal test involved instrumenting a flight-like valve with six microphones and then exciting the enclosed air with a loudspeaker. The loudspeaker was configured to deliver broadband noise at relatively high sound pressure levels. The aquired microphone signals were post-processed and compared to results generated from the acoustic finite element model. Initial comparisons between the test data and the model results revealed that additional model refinement was necessary. Specifically, the model was updated to implement a complex impedance boundary condition at the entrance to the valve supply tube. This boundary condition models the frequency-dependent impedance that an acoustic wave will encounter as it reaches the end of the supply tube. Upon invoking this boundary condition, significantly improved agreement between the test data and the model was realized.

Davis, R. Benjamin↗