Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Experimental Test Range”

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 937 records · Page 52
All sourcesNASA NTRSDOE OSTI
DOE OSTI2021

Machine learning elastic constants of multi-component alloys

The present manuscript explores application of machine learning methods for determining elastic constants and other derived mechanical properties of multi-component alloys. Here, a number of machine learning models, including linear regression, neural network and random forest based models, are trained and tested on a dataset of binary alloys generated using density functional theory (DFT) calculations and spanning over a large number of elemental species in the periodic table. Starting with a wide range of simple and easily accessible compositionally-averaged elemental features, a correlation-based feature selection strategy was used to systematically down-select a set of most relevant features towards the prediction of the elasticity tensor components. The true predictive performance and the associated uncertainties of the models were established by testing on unseen data and bootstrapping, respectively. A single and pair-wise feature partial dependence analysis was performed to visualize the average property trends in the multi-dimensional feature space in order to further understand the achieved predictive performance. The utility of the trained model is further demonstrated by obtaining sufficiently accurate yet highly efficient approximations for bulk modulus, Young’s modulus, shear modulus and Poisson’s ratio for alloys beyond the binary space (i.e., two-component alloys) on which the model was originally trained. More importantly, we test and validate the predictive performance of the developed model directly against the experimentally measured elastic constants of technologically relevant multi-component alloys (such as, Ni- and Ti-based alloys). Finally, utility of such a data-enabled route is demonstrated by predicting the possible range of various elastic properties for vast composition space available within the five component Ni-Cr-Fe-Mo-W alloy system in a high-throughput manner.

36 MATERIALS SCIENCE↗
NASA NTRS2015

Raven: An On-Orbit Relative Navigation Demonstration Using International Space Station Visiting Vehicles

Since the last Hubble Servicing Mission five years ago, the Satellite Servicing Capabilities Office (SSCO) at the NASA Goddard Space Flight Center (GSFC) has been focusing on maturing the technologies necessary to robotically service orbiting legacy assets-spacecraft not necessarily designed for in-flight service. Raven, SSCO's next orbital experiment to the International Space Station (ISS), is a real-time autonomous non-cooperative relative navigation system that will mature the estimation algorithms required for rendezvous and proximity operations for a satellite-servicing mission. Raven will fly as a hosted payload as part of the Space Test Program's STP-H5 mission, which will be mounted on an external ExPRESS Logistics Carrier (ELC) and will image the many visiting vehicles arriving and departing from the ISS as targets for observation. Raven will host multiple sensors: a visible camera with a variable field of view lens, a long-wave infrared camera, and a short-wave flash lidar. This sensor suite can be pointed via a two-axis gimbal to provide a wide field of regard to track the visiting vehicles as they make their approach. Various real-time vision processing algorithms will produce range, bearing, and six degree of freedom pose measurements that will be processed in a relative navigation filter to produce an optimal relative state estimate. In this overview paper, we will cover top-level requirements, experimental concept of operations, system design, and the status of Raven integration and test activities.

Rendezvous↗
DOE OSTI2023

Machine Learning for Prediction of Thermodynamic Descriptors

Our objective is to apply machine learning (ML) algorithms for the prediction of molecular catalysis descriptors from geometric properties derived from experimental crystallographic databases. Catalysis is often considered a “low-data” discipline that is poorly suited for ML methods. An exception is the extensive structural information that is available for molecular catalysts through the Cambridge Structural Database (CSD), which contains atomically precise molecular structures from X-ray diffraction analysis for >600K metal complexes. As a proof-of-principle, we targeted the prediction of hydricity, a thermodynamic property that provides understanding and control of catalytic hydride transfer. We built a training set composed of ~100 molecular complexes with a known hydricity and structural information from the CSD. This data set was converted into a machine-readable format using the smooth overlap of atomic positions (SOAP) representation and further labeled with simple electronic descriptors for the metal centers. Multiple different neural networks were trained on this data set, and the accuracy of the hydricity predictions ranged from < 2 kcal/mol to 20 kcal/mol. The accuracy of each model was highly sensitive to which compounds were in the train versus test set, underscoring the challenges associated with small and chemically diverse data sets. Finally, to further augment the data set, we attempted to experimentally measure several new hydricity values, however these experiments were unsuccessful due to undesired chemical reactivity of the selected complexes.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
NASA NTRS1986

An experimental investigation of the flap-lag-torsion aeroelastic stability of a small-scale hingeless helicopter rotor in hover

A small scale, 1.92 m diam, torsionally soft, hingeless helicopter rotor was investigated in hover to determine isolated rotor stability characteristics. The two-bladed, untwisted rotor was tested on a rigid test stand at tip speeds up to 101 m/sec. The rotor mode of interest is the lightly damped lead-lag mode. The dimensionless lead-lag frequency of the mode is approximately 1.5 at the highest tip speed. The hub was designed to allow variation in precone, blade droop, pitch control stiffness, and blade pitch angle. Measurements of modal frequency and damping were obtained for several combinations of these hub parameters at several values of rotor speed. Steady blade bending moments were also measured. The lead-lag damping measurements were found to agree well with theoretical predictions for low values of blade pitch angle. The test data confirmed the predicted effects of precone, droop, and pitch control stiffness parameters on lead-lag damping. The correlation between theory and experiment was found to be poor for the mid-to-high range of pitch angles where the theory substantially overpredicted the experimental lead-lag damping. The poor correlation in the mid-to-high blade pitch angle range is attributed to low Reynolds number nonlinear aerodynamics effects not included in the theory. The experimental results also revealed an asymmetry in lead-lag damping between positive and negative thrust conditions.

Sharpe, David L.↗
NASA NTRS1968

Propulsive re-entry aerodynamics Interim report, 1 Jun. 1967 - 1 Jun. 1968

This report discusses the progress accomplished during the first year of a two-year contract to study Propulsive Re-Entry Aerodynamics. Analytical methods were developed during the first year to predict the flow field in the immediate vicinity of a planetary entry system composed of an aeroshell with a single retrorocket which exhausts into a subsonic or supersonic counterflowing planetary atmosphere. In the second year a digital computer program will be developed for the analytical solutions, and the analytical model will be tested by performing an experimental test program and comparing the experimental results with the analytical predictions. The analytical investigation considers the flow field due to a single supersonic retrorocket exhausting from an aeroshell into oncoming subsonic, transonic, or supersonic streams in the large thrusting coefficient regime, 0.5 < C(sub T) = T/q(sub ∞) A(sub m) < 15, which is the range of thrusting coefficients encountered during the terminal landing phase for planetary landing on Mars or Venus. Analyses are provided for all the geometrical features of the flow field such as the location of the terminal shock, jet boundary, and the profile for the interface between the jet and atmospheric gases. In addition, analyses are developed which describe the growth of the mixing layers along the interface and at the edge of the jet boundary and dead-air region and the reattachment of the shear layer to the aeroshell or the recompression of the shear layer in wake type flow. A comprehensive discussion of the experimental wind tunnel test program to be accomplished in Phase II is presented.

Philip O Jarvinen↗
NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗
NASA NTRS1949

Hovering and Low-Speed Performance and Control Characteristics of the Kaman Helicopter Rotor System as Determined on the Langley Helicopter Tower. TED No. NACA DE 205

An investigation has been conducted with the Langley helicopter tower to obtain basic performance and control characteristics of the Raman rotor system. Blade-pitch control is obtained in this configuration by utilizing an auxiliary flap to twist the blades. Rotor thrust and power required were measured for the hovering condition and over a range of wind velocities from 0 to 30 miles per hour. The control characteristics and the transient response of the rotor to various control movements were also measured. The hovering-performance data are presented as a survey of the wake velocities and the variation of torque coefficient with thrust coefficient. The power required for the test rotor to hover at a thrust of 1350 pounds and a rotor speed of 240 rpm is approximately 6.5 percent greater than that estimated for a conventional rotor of the same diameter and solidity. It is believed that most of this difference is caused by th e flap servomechanism. The reduction in total power required for sustentation of the single-rotor configuration tested at various wind velocities and at the normal operating rotor thrust was found to be similar to the theoretical and experimental results for ro tors with conventionally actuated pitch. The control effectiveness was determined as a function of rotor speed. Sufficient control was available to give a thrust range of 0 to 1500 pounds and a rotor tilt of plus or minus 7 degrees. The time lag between flap motion and blade-pitch response is approximately 0.02 to 0.03 second. The response of the rotor following the blade-pitch response is similar to that of a rotor with conventionally actuated pitch changes. The over-all characteristics of the rotor investigated indicate that satisfactory performance and control characteristics were obtained.

Carpenter, Paul J.↗
NASA NTRS1972

Experimental study of flow distribution and pressure loss with circumferential inlet and outlet manifolds

Water flow tests with circumferential inlet and outlet manifolds were conducted to determine factors affecting fluid distribution and pressure losses. Various orifice sizes and manifold geometries were tested over a range of flow velocities. With inlet manifolds, flow distribution was related directly to orifice discharge coefficients. A correlation indicated that nonuniform distribution resulted when the velocity head ratio at the orifice was not in the range of constant discharge coefficient. With outlet manifolds, nonuniform flow was related to static pressure variations along the manifold. Outlet manifolds had appreciably greater pressure losses than comparable inlet manifolds.

Dittrich, R. T.↗
NASA NTRS2000

Engine-Scale Combustor Rig Designed, Fabricated, and Tested for Combustion Instability Control Research

Low-emission combustor designs are prone to combustor instabilities. Because active control of these instabilities may allow future combustors to meet both stringent emissions and performance requirements, an experimental combustor rig was developed for investigating methods of actively suppressing combustion instabilities. The experimental rig has features similar to a real engine combustor and exhibits instabilities representative of those in aircraft gas turbine engines. Experimental testing in the spring of 1999 demonstrated that the rig can be tuned to closely represent an instability observed in engine tests. Future plans are to develop and demonstrate combustion instability control using this experimental combustor rig. The NASA Glenn Research Center at Lewis Field is leading the Combustion Instability Control program to investigate methods for actively suppressing combustion instabilities. Under this program, a single-nozzle, liquid-fueled research combustor rig was designed, fabricated, and tested. The rig has many of the complexities of a real engine combustor, including an actual fuel nozzle and swirler, dilution cooling, and an effusion-cooled liner. Prior to designing the experimental rig, a survey of aircraft engine combustion instability experience identified an instability observed in a prototype engine as a suitable candidate for replication. The frequency of the instability was 525 Hz, with an amplitude of approximately 1.5-psi peak-to-peak at a burner pressure of 200 psia. The single-nozzle experimental combustor rig was designed to preserve subcomponent lengths, cross sectional area distribution, flow distribution, pressure-drop distribution, temperature distribution, and other factors previously found to be determinants of burner acoustic frequencies, mode shapes, gain, and damping. Analytical models were used to predict the acoustic resonances of both the engine combustor and proposed experiment. The analysis confirmed that the test rig configuration and engine configuration had similar longitudinal acoustic characteristics, increasing the likelihood that the engine instability would be replicated in the rig. Parametric analytical studies were performed to understand the influence of geometry and condition variations and to establish a combustion test plan. Cold-flow experiments verified that the design values of area and flow distributions were obtained. Combustion test results established the existence of a longitudinal combustion instability in the 500-Hz range with a measured amplitude approximating that observed in the engine. Modifications to the rig configuration during testing also showed the potential for injector independence. The research combustor rig was developed in partnership with Pratt & Whitney of West Palm Beach, Florida, and United Technologies Research Center of East Hartford, Connecticut. Experimental testing of the combustor rig took place at United Technologies Research Center.

DeLaat, John C.↗
DOE OSTI2021

Performance evaluation of a Terry GS-2 steam impulse turbine with air-water mixtures

Terry steam turbines are widely used in various industries because of their robust design. Within the nuclear power generation industry, they are used in the Reactor Core Isolation Cooling System to remove decay heat during reactor isolation events. During the Fukushima Daiichi nuclear power station disaster in Japan in 2011, the Reactor Core Isolation Cooling System and associated Terry turbine operated for over 70 hours in Unit 2; this runtime is well beyond the expected operating duration. Theories suggest the turbine was subjected to a two-phase inlet flow, which could degrade the turbine performance. In this work, an experimental test rig was constructed to test a full-scale Terry model GS-2 steam turbine under two-phase air/water flows. Steady-state efficiency and torque performance maps of the turbine were developed over a range of turbine inlet pressures (1.38–4.83 bar or 20–70 psia), air mass fractions (0.05–1.0) and rotational speeds up to 4000 RPM. Furthermore, turbine performance followed expected trends with torque varying linearly and efficiency varying quadratically with rotational speed. In addition, high-speed images of the two-phase flow entering the turbine were also analyzed to understand how changes in inlet pressure and air mass fraction affect the flow regime and homogenization. The present tests with air–water two-phase mixtures are an important step towards providing an understanding of the full-scale Terry turbine’s behavior and performance curves under two-phase conditions. The results of this work will be combined with air/water and steam/water data gathered using a small-scale Terry ZS-1 steam turbine in order to understand the scaling relationship between large and small size Terry turbines and fluid pairs. The combined data set will enable further development of analytical models over a wide range of conditions and may be used to provide technical justification for expanded use of the Terry turbines in nuclear power plant safety systems and other systems.

42 ENGINEERING↗
DOE OSTI2024

The Jefferson Lab Eta Factory Experiment and Applications of PbWO4 Calorimeters in Future Experimental Facilities

The goal of the new JLab Eta Factory (JEF) experiment, conducted with the GlueX detector in Hall D at Jefferson Lab, is to perform measurements of various ¿(') decays with a primary focus on rare neutral modes. The experiment’s physics program ranges from precision tests of low-energy QCD to searches for gauge bosons with masses below 1 GeV that could couple the Standard Model (SM) sector to the dark sector. The experiment will collect a high-statistics data sample of ¿(') mesons produced via a beam of tagged photons. The GlueX detector features a large, nearly uniform acceptance for both neutral and charged particles, enabling efficient identification of complex multi-particle final states. To meet the requirements of the JEF experiment, the inner section of the forward lead-glass calorimeter in the GlueX detector has been upgraded with lead tungstate (PbWO4) scintillating crystals. PbWO4 offers exceptional characteristics, such as a small radiation length and Molire radius, and large light yield, that make it ideal for constructing high- granularity, high-resolution, radiation-hard detectors. These properties enable excellent spatial separation and energy resolution of reconstructed electromagnetic showers, establishing PbWO4 as the material of choice for many high-precision experiments. The JEF experiment began data collection in April 2025 and will operate concurrently with the GlueX experiment, whose primary objective is the search for gluonic excitations in the meson spectrum. I will give an overview of the JEF experiment, the GlueX detector, and the feasibility of further upgrades to support future ¿ physics studies. Special attention will be given to the newly constructed PbWO4 scintillating calorimeter and recent advancements in calorimeter instrumentation.

Somov, Alexander↗
DOE OSTI2023

Snowmass 2021 Whitepaper: Proton Structure at the Precision Frontier

An overwhelming number of theoretical predictions for hadron colliders require parton distribution functions (PDFs), which are an important ingredient of theory infrastructure for the next generation of high-energy experiments. This whitepaper summarizes the status and future prospects for determination of high-precision PDFs applicable in a wide range of energies and experiments, in particular in precision tests of the Standard Model and in new physics searches at the high-luminosity Large Hadron Collider and Electron–Ion Collider. We discuss the envisioned advancements in experimental measurements, QCD theory, global analysis methodology, and computing that are necessary to bring unpolarized PDFs in the nucleon to the N2LO and N3LO accuracy in the QCD coupling strength. Special attention is given to the new tasks that emerge in the era of the precision PDF analysis, such as those focusing on the robust control of systematic factors both in experimental measurements and theoretical computations. Various synergies between experimental and theoretical studies of the hadron structure are explored, including opportunities for studying PDFs for nuclear and meson targets, PDFs with electroweak contributions or dependence on the transverse momentum, for incisive comparisons between phenomenological models for the PDFs and computations on discrete lattice, and for cross-fertilization with machine learning/AI approaches.

72 PHYSICS OF ELEMENTARY PARTICLES AND FIELDS↗
NASA NTRS1972

A search for the effect of a high frequency spark on boundary-layer transition at Mach 8.5

An experimental investigation of the use of a high frequency spark to promote early boundary layer transition on a wind tunnel model was conducted at a Mach number of 8.5. Test variables included four electrode configurations, a frequency range from 10 kHz to 50 kHz, and various power inputs to the spark. The general conclusion obtained from this investigation is that over the parameter range, the high frequency spark is ineffective in inducing early transition at the test Mach number.

Anders, J. B.↗
NASA NTRS1973

Design and experimental results for a turbine with jet flap stator and jet flap

The overall performance and detailed stator performance of a negative hub reaction turbine design featuring a moderately low solidity jet flap stator and a jet flap rotor were determined. Testing was conducted over a range of turbine expansion ratios at design speed. At each expansion ratio, the stator jet flow and rotor jet flow ranged up to about 7 and 8 percent, respectively, of the turbine inlet flow. The performance of the jet flap stator/jet flap rotor turbine was compared with that of a turbine which used the same jet flap rotor and a conventional, high solidity plan stator. The effect on performance of increased axial spacing between the jet stator and rotor was also investigated.

Bettner, J. L.↗
NASA NTRS1975

Aerodynamic performance of a ceramic-coated core turbine vane tested with cold air in a two-dimensional cascade

The aerodynamic performance of a core turbine vane with a ceramic coating was investigated experimentally in a two-dimensional cascade of 10 vanes. The vane was first tested with a rough coating and then retested after the coating had been smoothed. The test fluid was atmospheric air. The cascade tests covered a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95. The principal measurements were cross-channel surveys of exit total pressure, static pressure, and flow angle. The results presented include exit survey results and overall performance in terms of flow angle, weight flow, and kinetic energy loss. The performance of the ceramic-coated vanes is compared with the performance of an uncoated vane having a similar profile.

Stabe, R. G.↗
NASA NTRS1976

The U.S.-NSF/NASA

The five-year research and development plan of the NSF/NASA Wind Energy Conversion Systems (WECS) program is outlined. The program includes mission studies to determine energy use patterns and requirements and define specific applications for wind energy systems, wind energy resource assessment and development, and development of cost-effective components and subsystems. The program is also directed towards the development of energy storage systems to make wind powered systems firm power sources where appropriate. A 100 kW experimental wind generator (Model Zero) is being designed as a flexible test bed for a variety of system components. Designs will be developed for units in the 50 to 200 kW and 500 to 3000 kW size ranges.

Source record↗
NASA NTRS1984

Design and performance of a fixed, nonaccelerating, guide vane cascade that operates over an inlet flow angle range of 60 deg

A unique set of wind tunnel guide vanes are designed with an inverse design code and analyzed with a panel method and an integral boundary layer code developed at the NASA Lewis Research Center. The fixed guide vanes, 80 feet long with 6-foot chord length, were designed for the NASA Ames 40 x 80/80 x 120 ft Wind Tunnel. Low subsonic flow is accepted over a 60 deg range of inlet angle from either the 40 x 80 leg or the 80 x 120 leg of the wind tunnel, and directed axially into the main leg of the tunnel where drive fans are located. Experimental tests of 1/10-scale models were conducted to verify design calculations.

Sanz, J. M.↗
NASA NTRS1984

Shutdown characteristics of the Mod-O wind turbine with aileron controls

Horizontal-axis wind turbines utilize partial or full variable blade pitch to regulate rotor speed. The weight and costs of these systems indicated a need for alternate methods of rotor control. Aileron control is an alternative which has potential to meet this need. The NASA Lewis Research Center has been experimentally testing aileron control rotors on the Mod-U wind turbine to determine their power regulation and shutdown characteristics. Experimental and analytical shutdown test results are presented for a 38 percent chord aileron-control rotor. These results indicated that the 38 percent chord ailerons provided overspeed protection over the entire Mod-O operational windspeed range, and had a no-load equilibrium tip speed ratio of 1.9. Thus, the 38 percent chord ailerons had much improved aerodynamic braking capability when compared with the first aileron-control rotor having 20 percent chord ailerons.

Miller, D. R.↗