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At least 163 records · Page 9

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NASA NTRS1981

Fumigation of Alcohol in a Light Duty Automotive Diesel Engine

A light-duty automotive Diesel engine was fumigated with methanol in amounts up to 35% and 50% of the total fuel energy respectively in order to determine the effect of alcohol fumigation on engine performance at various operating conditons. Engine fuel efficiency, emissions, smoke, and the occurrence of severe knock were the parameters used to evaluate performance. Raw exhaust particulate and its soluble organic extract were screened for biological activity using the Ames Salmonella typhimurium assay. Results are given for a test matrix made up of twelve steady-state operating conditions. For all conditions except the 1/4 rack (light load) condition, modest thermal efficiency gains were noted upon ethanol fumigation. Methanol showed the same increase at 3/4 and full rack (high load) conditions. However, engine roughness or the occurrence of severe knock limited the maximum amount of alcohol that could be fumigated. Brake specific nitrogen oxide concentrations were found to decrease for all ethanol conditions tested. Oxides of nitrogen emissions, on a volume basis, decreased for all alcohol conditions tested. Based on the limited particulate data analyzed, it appears that ethanol fumigation, like methanol fumigation, while lowering the mass of particulated emitted, does enhance the biological activity of that particulate.

Broukhiyan, E. M. H.↗

NASA NTRS2011

Overview of Experimental Investigations for Ares I Launch Vehicle Development

Another concern for the vehicle during its design trajectory was the separation of the first stage solid rocket booster from the upper stage component after it had depleted its solid fuel propellant. There has been some concern about the interstage of the first stage from clearing the nozzle of the J2-X engine. A detailed separation aerodynamic wind tunnel investigation was conducted in the AEDC VKF Tunnel A to help to investigate the interaction aerodynamic effects5. A comparison of the separation plane details between the Ares I architecture and the Ares I-X demonstration flight architecture is shown in figure 12. The Ares I design requires a more complex separation sequence and requires better control in order to avoid contact with the nozzle of the upper stage engine. The interstage, which houses the J2-X engine for the Ares I vehicle, must be able to separate cleanly to avoid contact of the J2-X engine. There is only about approximately 18 inches of buffer inside the interstage on each size of the nozzle so this is a challenging controlled separation event. This complex experimental investigation required two separate Ares I models (upper stage and first stage with interstage attached) with independent strain gauge balances installed in each model. It also required the Captive Trajectory System (CTS) that was needed to precisely locate the components in space relative to each other to fill out the planned test matrix. The model setup in the AEDC VKF Tunnel A is shown in figure 13. The CTS remotely positioned the first stage at the required x, y, and z positions and was able to provide interactions within 0.2" of the upper stage. A sample of the axial force on the first stage booster is shown in figure 14. These results, as a function of separation distance between the two stages, are compared to pre-test CFD results. Since this is a very challenging, highly unsteady flow field for CFD to correctly model, the experimental results have been utilized by GN&C discipline to more accurately represent the interaction aerodynamics. In addition to the integrated forces and moments obtained from the test, flow visualization data was obtained from this test in the form of Schlieren photographs, as shown in figure 15, which show the shock structure and interaction effects after the two stages separate during flight. This separation test was crucial in the successful flight test of the Ares I-X vehicle and provided the GN&C discipline with the unpowered proximity aerodynamic effect for a separation of the Ares I vehicle.

Tomek, William G.↗

NASA NTRS2012

Acoustic Surveys of a Scaled-Model CESTOL Transport Aircraft in Static and Forward Speed Conditions

An 11% scale-model of a Cruise-Efficient Short Take-off and Landing (CESTOL) scalemodel test was recently completed. The test was conducted in the AEDC National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center. The model included two over-wing pod-mounted turbine propulsion simulators (TPS). The hybrid blended wing-body used a circulation control wing (CCW) with leadingand trailing-edge blowing. The bulk of the test matrix included three forward velocities (40 kts, 60 kts, and 100kts), angle-of-attack variation between -5 and 25 , and CCW mass flow variation. Seven strut-mounted microphones outboard of the left wing provided source directivity. A phased microphone array was mounted outboard of the right wing for source location. The goal of this paper is to provide a preliminary look at the acoustic data acquired during the Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA) test for 0 angle-of-attack and 0 sideslip conditions. Data presented provides a good overview of the test conditions and the signal-to-noise quality of the data. TPS height variation showed a difference of 2 dB to 3 dB due to wing shielding. Variation of slot mass flow showed increases of 12 dB to 26 dB above the airframe noise and the TPS increased the overall levels an additional 5 dB to 10 dB.

Burnside, Nathan↗

NASA NTRS2020

DGEN Aeropropulsion Research Turbofan Source-Diagnostic Test: Experimental Setup and Acoustic-Data Structure

The experimental setup of, and available data from a recent core/combustor-noise source-diagnostic test utilizing a small turbofan engine are described. The 2019 test campaign continued the investigation of the core/combustor-noise component of aircraft-propulsor noise begun in an earlier baseline test, but with a more extensive acoustic-instrumentation layout. The purpose of both tests was to better understand the impact on civilian-transport airport-community noise from turbofan-combustor sources and thereby to lay the foundation for improved noise-prediction methods and noise-mitigation techniques. Simultaneous high-data-rate acoustic measurements were obtained using a circumferential sensor array at the core nozzle exit in conjunction with sideline and farfield microphone arrays. The test matrix contained engine operational points from engine idle to maximum power and was repeated for different circumferential and sideline array configurations, as well as for redundancy. The extensive data set (up to 93 channels of data and various configurations) allows the application of advanced source-separation and phased-array methods to elucidate not only the core-noise structure, but also the propagation characteristics of other propulsion noise sources. The present report provides a detailed description of the different test points, their associated instrumentation layouts, and the structure of the acquired data set. Results from various data analyses are reported separately.

Aeroacoustics↗

NASA NTRS1985

Test results of a 40 kW Stirling engine and comparison with the NASA-Lewis computer code predictions

A Stirling engine was tested without auxiliaries at NASA-Lewis. Three different regenerator configurations were tested with hydrogen. The test objectives were (1) to obtain steady-state and dynamic engine data, including indicated power, for validation of an existing computer model for this engine; and (2) to evaluate structurally the use of silicon carbide regenerators. This paper presents comparisons of the measured brake performance, indicated mean effective pressure, and cyclic pressure variations with those predicted by the code. The measured data tended to be lower than the computer code predictions. The silicon carbide foam regenerators appear to be structurally suitable, but the foam matrix tested severely reduced performance.

Allen, D.↗

NASA NTRS2020

PIV and Rotational Raman-Based Temperature Measurements for CFD Validation of a Perforated Plate Cooling Flow: Part I

Film cooling is used in a wide variety of engineering applications for protection of surfaces from hot or combusting gases. The design of more efficient film cooling geometries/configurations could be facilitated by an ability to accurately model and predict the effectiveness of current designs using computational fluid dynamics (CFD) code predictions. Hence, a benchmark set of flow field property data was obtained to assess current CFD capabilities and develop better modeling approaches for these turbulent flow fields where accurate calculation of turbulent heat flux is important. Both Particle Image Velocimetry (PIV) and spontaneous rotational Raman scattering (SRS) spectroscopy were used to acquire high quality, spatially-resolved measurements of the mean and root mean square (rms) velocities as well as the mean and rms temperatures in a film cooling flow field. In addition to off-body flow field measurements, infrared thermography (IR) and thermocouple measurements on the plate surface enabled estimates of the film effectiveness. Raman spectra in air were obtained across a matrix of axial locations downstream from a 68.07 mm square nozzle blowing heated air over a range of temperatures (up to TR = 2.7) and Mach numbers (up to Mach 0.9), across a 30.48 cm long plate equipped with three patches of 45 small (~1 mm) diameter cooling holes arranged in a staggered configuration. In addition, both streamwise 2-component PIV (on the plate centerline) and cross-stream 3-component Stereo PIV data at 14 axial stations were collected in the same flows. Only a subset of the data collected in the test program is included in this Part I report. The rest of the data will be published in a future report, Part II, along with planned CFD predictions of the complex cooling film flow. The entire data set of Raman temperature data, PIV velocity data and IR camera data covering the Set Points 23 and 49 in the test matrix in Table 1 is available in an accompanying DVD (available online from www.sti.nasa.gov) for those interested in further analysis.

Nozzle↗

NASA NTRS2020

PIV and Rotational Raman-Based Temperature Measurements for CFD Validation of a Perforated Plate Cooling Flow Part I

Film cooling is used in a wide variety of engineering applications for protection of surfaces from hotor combusting gases. The design of more efficient film cooling geometries/configurations could be facilitated by an ability to accurately model and predict the effectiveness of current designs using computational fluid dynamics (CFD) code predictions. Hence, a benchmark set of flow field property data were obtained to assess current CFD capabilities and develop better modeling approaches for these turbulent flow fields where accurate calculation of turbulent heat flux is important. Both Particle Image Velocimetry (PIV) and spontaneous rotational Raman scattering (SRS) spectroscopy were used to acquire high quality, spatially-resolved measurements of the mean and root mean square (rms) velocities as well as the mean and rms temperatures in a film cooling flow field. In addition to off-body flow field measurements, infrared thermography (IR) and thermocouple measurements on the plate surface enabled estimates of the film effectiveness. Raman spectra in air were obtained across a matrix of axial locations downstream from a 68.07 mm square nozzle blowing heated air over a range of temperatures (up to TR =2.7) and Mach numbers (up to Mach 0.9), across a 30.48 cm long plate equipped with three patches of 45 small (~1 mm) diameter cooling holes arranged in a staggered configuration. In addition, both streamwise 2-component PIV (along the plate centerline) and cross-stream 3-component Stereo PIV data at 14 axial stations were collected in the same flows. Only a subset of the data collected in the test program is included in this Part 1 report. The rest of the data will be published in a future report, Part 2, along with planned CFD predictions of the complex cooling film flow.The entire data set of Raman temperature data, PIV velocity data and IR camera data covering the Set Points 23 and 49 in the test matrix in Table 1 is available in an accompanying DVD (available online from www.sti.nasa.gov) for those interested in further analysis.

Mark P Wernet↗

NASA NTRSDate not supplied

Exploring Airfoil Table Generation using XFOIL and OVERFLOW

The rotorcraft design process is a continuously evolving field of research that incorporates a number of software programs. An accurate airfoil table is critical in the design and testing process for rotorcraft. With multiple flow solvers available and flow conditions of multirotor UAM vehicles potentially covering a wide range of Reynolds and Mach numbers, a documented approach for developing airfoil tables is needed. Using benchmark data from legacy airfoil tables and wind tunnel tests for comparison, simulations for a comprehensive test matrix could guide rotorcraft design engineers in generating their own airfoil tables using the XFOIL and OVERFLOW solvers. The motivation for this study is to investigate flow solver features to develop a best practices document for airfoil table generation. The study uses the OVERFLOW and XFOIL flow solvers, coupled with the airfoil table generator AFTGen, to analyze three airfoils for a specific Reynolds numbers flow regime and provide details on how well each flow solver performs within a specific angle of attack range, Mach number range, Reynolds number range, and in different flow conditions, such as turbulent and transitional flow. OVERFLOW analyses in AFTGen for fully turbulent and transition flow are compared with XFOIL results and experimental test data for the section lift, section drag, and pressure coefficients. XFOIL ultimately yields results that are accurate within the linear angle of attack range and below a Mach number of 0.4 but tends to overpredict lift and underpredict drag unless the flow is in the compressible regime. XFOIL cannot accurately model stall and post-stall conditions due to the nature of the solver. This is evident in nearly every case run with XFOIL, where the linear range is usually predicted acceptably and the lift coefficient is overpredicted as the stall angle of attack is approached (with the exception being the generally poor correlation with most of the SSC-A09 cases). OVERFLOW is limited at low Mach numbers, and appears to perform best at Mach numbers of 0.4 and above. The exploration of airfoil table generation using XFOIL and OVERFLOW yielded moderately successful results for the NACA 0012 airfoil, reasonably good results for the RC(4)-10 airfoil, and less accurate results for the SSC-A09 airfoil.

Airfoil Table Generation↗

NASA NTRSDate not supplied

Experimental Characterization of an Electric Ducted Fan for the SUSAN 25% Scale Research Vehicle

An electric ducted fan (EDF) was tested in the NASA Langley 12-Foot Low-Speed Tunnel for the SUbsonic Single Aft eNgine (SUSAN) Electrofan 25% scale research vehicle, which incorporates distributed electric fans along the wings and serves as a scaled electrified transport aircraft testbed. The commercial-off-the-shelf EDF was selected as a propulsor that can meet the 25% scale vehicle geometry constraints, along with the performance requirements throughout the flight profile. The test included a series of one-factor-at-a-time (OFAT) and design of experiments (DOE) approaches that were implemented across a range of fan speed, voltage, freestream dynamic pressure, and incidence angle settings to characterize the performance of the EDF. Several test matrix designs and modeling approaches were considered. System identification methods were used to develop a static motor model, a dynamic motor model, an EDF thrust model, and a power consumption model. The model predictive capabilities were then evaluated with validation data not used in the modeling process. The results provide a comprehensive set of models developed to characterize the fan, and offer general insight into EDF modeling approaches.

electric ducted fan↗

NASA NTRSDate not supplied

Preliminary Findings of the Experimental Development Unit Cold Flow Test for a Generation Zero Nuclear Propulsion Engine

Nuclear thermal propulsion (NTP) technology will greatly benefit human travel to Mars by significantly shortening transit times, improving crew safety, and providing more mission flexibility than traditional chemical rockets. As part of DRACO follow-on work to develop, build, and fly a generation zero NTP engine, a full scale flight-like experimental design unit (EDU) reactor was constructed to collect sufficient on-ground performance data to characterize flow induced vibrations (FIV) of critical reactor structures/components, inform development of the engine and reactor control algorithm, and collect pressure drop and flow distribution data across the reactor. The fluid conditions for the test program were designed to achieve system responses equivalent to that of an operational engine through all phases of engine operation including reactor startup, mainstage operation, reactor shutdown, and reactor cooldown. Over 100 tests were executed, flowing either GN2 or GHe through the EDU at varying flow rates and pressures. This experiment provided early validation of flow behavior and vibration risks before nuclear testing, boosted critical subsystem TRLs, informed design iterations, and reduced future test costs. The steady-state flow parameters for the experiment were modeled in Ansys Thermal Desktop, allowing rapid tuning and experiment-informed updates to a flight-like test matrix. The EDU dynamic environment was characterized with accelerometers, strain gauges, and high-frequency pressure transducers all sampled at 20 kHz. While many narrow-band oscillations were identified, no significant FIV occurred; the reactor structural responses tend to be enveloped by typical launch vehicle ascent vibration environments (defined up to 2 kHz), although significant energy is also present at higher frequencies.

Flow Induced Vibration↗

NASA NTRS2004

Probabilistic Residual Strength Model Developed for Life Prediction of Ceramic Matrix Composites

For the next generation of reusable launch vehicles, NASA is investigating introducing ceramic matrix composites (CMCs) in place of current superalloys for structural propulsion applications (e.g., nozzles, vanes, combustors, and heat exchangers). The higher use temperatures of CMCs will reduce vehicle weight by eliminating and/or reducing cooling system requirements. The increased strength-to-weight ratio of CMCs relative to superalloys further enhances their weight savings potential. However, in order to provide safe designs for components made of these new materials, a comprehensive life prediction methodology for CMC structures needs to be developed. A robust methodology for lifing composite structures has yet to be adopted by the engineering community. Current industry design practice continues to utilize deterministic empirically based models borrowed from metals design for predicting material life capabilities. The deterministic nature of these models inadequately addresses the stochastic character of brittle composites, and their empirical reliance makes predictions beyond the experimental test conditions a risky extrapolation. A team of engineers at the NASA Glenn Research Center has been developing a new life prediction engineering model. The Probabilistic Residual Strength (PRS) model uses the residual strength of the composite as its damage metric. Expected life and material strength are both considered probabilistically to account for the observed stochastic material response. Extensive experimental testing has been carried out on C/SiC (a candidate aerospace CMC material system) in a controlled 1000 ppm O2/argon environment at elevated temperatures of 800 and 1200 C. The test matrix was established to allow observation of the material behavior, characterization of the model, and validation of the model's predictive capabilities. Sample results of the validation study are illustrated in the graphs.

Thomas, David J.↗

NASA NTRS1989

Characterization testing of a 40 Ahr bipolar nickel hydrogen battery

In a continuing effort to develop NiH2 bipolar technology to a point where it can be used efficiently in space flight, testing of a second 40 Ahr, 10-cell bipolar battery has begun. This battery has undergone extensive characterization testing to determine the effects of such operating parameters as charge and discharge rates, temperature, and pressure. The fundamental design of this actively cooled bipolar battery is the same as the first battery. Most of the individual components, however, are from different manufacturers. Different testing procedures as well as certain unique battery characteristics make it difficult to directly compare the two sets of results. In general, the performance of this battery throughout characterization produced expected results. The main differences seen between the first and second batteries occurred during the high-rate discharge portion of the test matrix. The first battery also had poor high-rate discharge results, although better than those of the second battery. Minor changes were made to the battery frame design used for the first battery in an attempt to allow better gas access to the reaction sites for the second build and hopefully improve performance. The changes, however, did not improve the performance of the second battery and could have possibly contributed to the poorer performance that was observed. There are other component differences that could have contributed to the poorer performance of the second battery. The H2 electrode in the second battery was constructed with a Goretex backing which could have limited the high-rate current flow. The gas screen in the second battery had a larger mesh which again could have limited the high-rate current flow. Small scale 2 x 2 batteries are being tested to evaluate the effects of the component variations.

Brewer, Jeffrey C.↗

NASA NTRS2024

Core/Combustor-Noise Measurements and Source Separation for the DGEN Aeropropulsion Research Turbofan

Acoustic data obtained using a small turbofan engine are further analyzed with the aim to ultimately enable improved prediction methods and mitigation techniques for turbofan core/combustor noise. The relative impact of this propulsion-noise source for civilian-transport aircraft on airport community noise is expected to significantly increase in the future. Simultaneous high-data-rate acoustic measurements acquired using a circumferential infinite-tube-probe array at the core-nozzle exit in conjunction with a far-field microphone array are processed. The test matrix contains engine operational points from idle to engine-maximum setting, with measurements repeated for different circumferential array configurations, as well as for redundancy. The combustor-noise contribution to the far-field noise signature is obtained using an advanced source-separation method that correlates far-field microphone measurements with a modal decomposition of the unsteady pressure field at the core-nozzle exit. The advantages of the present approach compared to the classical two-signal source-separation method are discussed.

Aeroacoustics↗

NASA NTRS2021

Characterization of Chevron Nozzle Performance

The flow fields from a set of 50.8 mm diameter circular nozzles were characterized using both streamwise and cross-stream 3-component Stereo Particle Image Velocimetry (SPIV). The test matrix of nozzles included 15 different chevron nozzle designs and one baseline, circularly symmetric nozzle. Nozzle jet operating conditions ranged from acoustic Mach numbers of 0.9 to 1.5, with static temperature ratios ranging between 0.84 and 2.7. Detailed surveys of the single jet flows were performed to capture three-dimensional features of the turbulent exhaust jet evolution. Cross-flow planar measurements were obtained at twelve axial locations, ranging from 0.1 to 20 nozzle diameters downstream of the nozzle exit planes. Streamwise measurements, along the jet centerlines, were obtained at ten partially overlapping downstream locations, providing complete axial surveys over a region extending beyond 20 nozzle diameters downstream of the nozzle exit planes. In both optical configurations, the measurement planes were sized to completely capture the fully turbulent jet shear layer growth. The measured three-dimensional mean and turbulent velocity fields, along with computed second order statistics including axial vorticity and turbulent kinetic energy, were evaluated for all test points. Well-defined streamwise vortex structures in the jet shear layers were measured and reported.

Jet Noise↗

NASA NTRS2016

A Comparison of the Noise Characteristics of a Conventional Slat and Krueger Flap

An aeroacoustic test of two types of leading-edge high-lift devices has been conducted in the NASA Langley Quiet Flow Facility. The test compares a conventional slat with a notional equivalent-mission Krueger flap. The test matrix includes points that allow for direct comparison of the conventional and Krueger devices for equivalent-mission configurations, where the two high-lift devices satisfy the same lift requirements for a free air flight path at the same cruise airfoil angle of attack. Measurements are made for multiple Mach numbers and directivity angles. Results indicate that the Krueger flap shows similar agreement to the expected power law scaling of a conventional flap, both in terms of Strouhal number and fixed frequency (as a surrogate for Helmholtz number). Directivity patterns vary depending on the specific slat and Krueger orientations. Varying the slat gap while holding overlap constant has the same influence on both the conventional slat and Krueger flap acoustic signature. Closing the gap shows dramatic reduction in levels for both devices. Varying the Krueger overlap has a different effect on the data when compared to varying the slat overlap, but analysis is limited by acoustic sources that regularly present themselves in model-scale wind tunnel testing but are not present for full-scale vehicles. The Krueger cavity is found to have some influence on level and directivity, though not as much as the other considered parameter variations. Overall, while the spectra of the two devices are different in detail, their scaling behavior for varying parameters is extremely similar.

Bahr, Christopher J.↗

NASA NTRS2016

Analysis and Characterization of Damage Utilizing an Orthotropic Generalized Composite Material Model Suitable for Use in Impact Problems

The need for accurate material models to simulate the deformation, damage and failure of polymer matrix composites under impact conditions is becoming critical as these materials are gaining increased usage in the aerospace and automotive communities. In order to address a series of issues identified by the aerospace community as being desirable to include in a next generation composite impact model, an orthotropic, macroscopic constitutive model incorporating both plasticity and damage suitable for implementation within the commercial LS-DYNA computer code is being developed. The plasticity model is based on extending the Tsai-Wu composite failure model into a strain hardening-based orthotropic plasticity model with a non-associative flow rule. The evolution of the yield surface is determined based on tabulated stress-strain curves in the various normal and shear directions and is tracked using the effective plastic strain. To compute the evolution of damage, a strain equivalent semi-coupled formulation is used in which a load in one direction results in a stiffness reduction in multiple material coordinate directions. A detailed analysis is carried out to ensure that the strain equivalence assumption is appropriate for the derived plasticity and damage formulations that are employed in the current model. Procedures to develop the appropriate input curves for the damage model are presented and the process required to develop an appropriate characterization test matrix is discussed.

Impact↗

NASA NTRS2016

Analysis and Characterization of Damage Utilizing an Orthotropic Generalized Composite Material Model Suitable for Use in Impact Problems

Impact↗

NASA NTRS1970

A frequency domain approach to handling qualities design

Linear multivariable feedback control systems design method based on transfer matrix, testing decoupling desirability

Shirley, R. S.↗