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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.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1953

Performance of a turbojet engine with adjustable first-stage turbine stator and variable-area exhaust nozzle

The performance of a turbojet engine with a two-stage turbine, an adjustable first-stage turbine stator, and a variable-area exhaust nozzle was investigated at selected constant engine speeds and two simulated flight conditions. For the particular component characteristics of the engine investigated, little improvement in thrust levels of interest by use of an adjustable rather than an optimum fixed first-stage turbine stator.

ENGINES, TURBOJET↗
NASA NTRS1953

Preliminary Evaluation of Turbine Performance with Variable-Area Turbine Nozzles in a Turbojet Engine

The performance of a two-stage turbine with variable-area first-stage turbine nozzles was determined in the NACA Lewis altitude wind tunnel over a range of simulated altitudes from 15,000 to 44,000 feet and engine speeds from 50 to 100 percent of rated speed. The variable-area turbine nozzles used in this investigation were primarily a test device for compressor research purposes and were not necessarily of optimum aerodynamic design. The results of this investigation are indicative of effects of turbine-nozzle-area variation on turbine performance within the operating range allowed by the engine. The variable-area turbine nozzles were found to be mechanically reliable and to have negligible leakage losses. Increasing the turbine-nozzle-throat area from 1.15 to 1.67 square feet increased the corrected turbine gas flow or effective turbine nozzle area about 10 percent. At a given corrected turbine speed and turbine pressure ratio, changing the turbine nozzle area from 1.30 to 1. 67 square feet lowered the turbine efficiency 3 or 4 percent. The effect of increasing the turbine nozzle area from 1.15 to 1.67 square feet (decreasing the turning angle about 7 1/2 degrees) would be to lower the turbine efficiency about 5 or 6 percent.

Campbell, Carl E.↗
NASA NTRS1957

Altitude Performance of a Turbojet Engine Using Pentaborane Fuel

A full-scale turbojet engine having a two-stage turbine was operated with pentaborane fuel continuously for 11.5 minutes at a simulated altitude of 55,000 feet at a flight Mach number of 0.8. The engine incorporated an NACA combustor designed specifically for use with pentaborane fuel. The specific fuel consumption was initially reduced 32 percent below that obtained with gasoline fuel; however, the occurrence of a 25-percent reduction in net thrust after 8 minutes of operation resulted in a subsequent increase in specific fuel consumption to a value only 11.5 percent lower than that for gasoline.

FUELS-RELATION TO ENGINE PERFORMANCE↗
NASA NTRS1947

Effect of Modifications to Induction System on Altitude Performance of V-1710-93 Engine III : Use of Parabolic Rotating Guide Vanes and NACA Designed Auxiliary-stage Inlet Elbow and Interstage Duct

Bench runs of a modified V-1710-93 engine equipped with a two-stage supercharger, interstage carburetor, aftercooler assembly, and backfire screens have been made at a simulated altitude of 29,000 feet to determine the effect of several induction-system modifications on the engine and supercharger performance. The standard guide vanes on the auxiliary- and engine-stage superchargers were replaced by rotating guide vanes with a parabolic blade profile. The auxiliary-stage inlet elbow and interstage duct were replaced with new units of NACA design. These modifications were made one at a time and data were obtained after each change to determine the effect of each modification. All runs were made at a constant engine speed of 3000 rpm at a simulated altitude of 29,000 feet and all changes in engine power were made by varying the speed of the auxiliary-stage supercharger.

Standahar, Ray M↗
NASA NTRS1996

Hypervelocity Impact Testing of Nickel Hydrogen Battery Cells

Nickel-Hydrogen (Ni/H2) battery cells have been used on several satellites and are planned for use on the International Space Station. In January 1992, the NASA Lewis Research Center (LeRC) conducted hypervelocity impact testing on Ni/H2 cells to characterize their failure modes. The cell's outer construction was a 24 mil-thick Inconel 718 pressure vessel. A sheet of 1.27 cm thick honeycomb was placed in front of the battery cells during testing to simulate the on-orbit box enclosure. Testing was conducted at the NASA White Sands Test Facility (WSTF). The hypervelocity gun used was a 7.6 mm (0.30 caliber) two-stage light gas gun. Test were performed at speeds of 3, 6, and 7 km/sec using aluminum 2017 spherical particles of either 4.8 or 6.4 mm diameter as the projectile. The battery cells were electrically charged to about 75 percent of capacity, then back-filled with hydrogen gas to 900 psi simulating the full charge condition. High speed film at 10,000 frames/sec was taken of the impacts. Impacts in the dome area (top) and the electrode area (middle) of the battery cells were investigated. Five tests on battery cells were performed. The results revealed that in all of the test conditions investigated, the battery cells simply vented their hydrogen gas and some electrolyte, but did not burst or generate any large debris fragments.

Frate, David T.↗
NASA NTRS1959

1958 NASA/USAF Space Probes (Able-1): Summary - Volume 1

Early in calendar year 1958 Space Technology Laboratories, Inc. (STL) (then Space Technology Laboratories, a division of the Ramo-Wooldridge Corp.) developed for the Air Force Ballistic Missile Division (AFBMD) an Advanced Re-entry Test Vehicle (ARTV) for the purpose of testing ballistic missile nose cones at the full range of 5500 nautical miles. The two-stage ARTV utilized the Thor ballistic missile and the second stage propulsion system developed for the Vanguard program. In late 1957 and early 1958, STL/AFBMD prepared studies of various missile combinations which could be utilized for space testing. The Thor, in combination with the Vanguard second and third stages, was one of the vehicles considered which offered a very early capability of placing a reasonable payload in a lunar orbit. These STL/AFBMD studies were presented to various appropriate groups including the Killian, Millikan, H. J . Stewart Committees; Headquarters, Air Research and Development Command, and ARDC Centers. Subsequently the Advanced Research Projects Agency (ARPA) contacted STL relative to the availability of hardware for an early lunar shot. By utilizing existing spares already purchased for the ARTV, and by making use of the ARTV contractors already in being, it appeared feasible to launch by the third quarter of calendar year 1958 a payload which would be captured by the moon's gravitational force. On 27 March 1958, ARPA directed STL to proceed with a program of three lunar shots. As much as possible, these shots were to utilize existing ARTV spare hardware and impose no interference with the ballistic missile programs. In September this program was transferred to the direction of the National Aeronautics and Space Administration (NASA). On 17 August 1958 the first launching of the Able-1 vehicle was attempted, but the flight was terminated by a propulsion failure of the first stage. Subsequent launchings were attempted on 13 October and 8 November 1958. Of these launchirigs the October attempt was the most successful. Although the payload did not reach the vicinity of the moon, a maximum altitude of 71,700 was attained, and useful scientific data was obtained from the instrumentation.

Source record↗
NASA NTRS1959

Free-Flight Investigation of a Rocket-Propelled Model to Determine the Aerodynamic Heating on a Thin, Unswept, Untapered, Multispar, Aluminum-Alloy Wing at Mach Numbers up to 2.22

A free-flight investigation has been made to determine some effects of aerodynamic heating on the structural behavior of a wing at supersonic speeds. The test wing was a thin, unswept, untapered, multispar, aluminum-alloy wing having a 20-inch chord, a 20-inch exposed semispan, and a circular-arc airfoil section with a thickness ratio of 5 percent. The wing was tested on a model propelled by a two-stage rocket-propulsion system to a Mach number of 2.22 and a corresponding Reynolds number per foot of 13.2 x 10(6) Reasonably good agreement was obtained between Stanton numbers obtained from measured temperature-time data and values obtained by the theory of Van Driest for flat plates having turbulent boundary layers. Temperature measurements made in the skin of the wing and in the internal structures agreed well with calculated values. The wing was instrumented to detect any apparent fluttering motion in the wing, but no evidence of flutter was observed throughout the flight.

Stephens, Emily W.↗
NASA NTRS1947

Investigation of Turbine of Mark 25 Torpedo Power Plant with Five Nozzle Designs

Efficiency investigations were made on the two-stage turbine from a Mark 25 aerial torpedo to determine the performance of the unit with five different turbine nozzles. The output of the turbine blades was computed by analyzing the windage and mechanical-friction losses of the unit. A method was developed for measuring the change in turbine clearances with changed operating conditions. The turbine was found to be most efficient with a cast nozzle having a sharp-edged inlet to the nine nozzle ports.

Hoyt, Jack W.↗
NASA NTRS1948

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine: Analysis of Effects of Inlet Pressure Losses - 6

The losses in the inlet air ducts, the diffusers, and the de-icing equipment associated with turbojet engine installations cause a reduction in the total pressure at the inlet of the engine and result in reduced thrust and increased specific fuel consumption. An analytical evaluation of the effects of inlet losses on the net thrust and the fuel economy of a 3000-pound-thrust axial flow turbojet engine with a two-stage turbine is presented. The analysis is based on engine performance characteristics that were determined from experiments in the NACA Cleveland altitude wind tunnel. The experimental investigation did not include tests in which inlet losses were systematically varied, but the effects of these losses can be accurately estimated from the experimentally determined performance characteristics of the engine.

Sanders, Newell D.↗
NASA NTRS1950

Summary of Investigations of Mark 25 Aerial-Torpedo Turbine

The power plant from a Mark 25 aerial torpedo was investigated both as a two-stage turbine and as a single-stage modified turbine to determine the effect on overall performance of nozzle size and shape, first-stage rotor-blade configuration, and axial nozzle-rotor running clearance. Performance was evaluated in terms of brake, rotor, and blade efficiencies. All the performance data were obtained for inlet total to outlet static pressure ratios of 8, 15 (design), and 20 with inlet conditions maintained constant at 95 pounds per square inch gage and 1000 F for rotor speeds from approximately 6000 to 18,000 rpm.

Schum, Harold J.↗
NASA NTRS1941

Performance Characteristics of an Aircraft Engine with Exhaust Turbine Supercharger, Special Report

The Pratt and Whitney Aircraft company and the Naval Aircraft Factory of the United States Navy cooperated in a laboratory and flight program of tests on an exhaust turbine supercharger. Two series of dynamometer tests of the engine super-charger combination were completed under simulated altitude conditions. One series of hot gas-chamber tests was conducted by the manufacturer of the supercharger. Flight demonstrations of the supercharger installed in a twin-engine flying boat were terminated by failure of the turbine wheels. The analysis of the results indicated that a two-stage supercharger with the first-stage exhaust turbine driven will deliver rated power for a given indicated power to a higher altitude, will operate more efficiently, and will require simpler controls than a similar engine with the first stage of the supercharger driven from the crankshaft through multispeed gears.

Lester, E. M.↗
NASA NTRS1952

Effect of Various Blade Modifications on Performance of a 16-Stage Axial-Flow Compressor. II - Effect on Over-All Performance Characteristics of Increasing Twelfth through Fifteenth Stage Stator-Blade Angles 3 deg

The stator-blade angles in the twelfth through fifteenth stages of a 16-stage axial-flow compressor were increased 3O. The over-all performance of this modified compressor is compared to the performance of the compressor with original blade angles. The matching characteristics of the modified compressor and a two-stage turbine were obtained and compared to those of the compressor with original blade angles and the same turbine.

Hatch, James E.↗
NASA NTRS1952

Effect of Various Blade Modifications on Performance of a 16-Stage High-pressure-ratio Axial-flow Compressor.Angles 3 deg: Effect on Over-all Performance Characteristics of Decreasing Twelfth Through Fifteenth State Stator-blade Angles 3 deg - 1

The stator-blade angles in the twelfth to fifteenth stages of a 16-stage high-pressure-ratio axial-flow compressor were decreased 3 deg The over-all performance of this compressor is compared with the performance of the same compressor with standard blade angles. The matching characteristics of the modified compressor and a two-stage turbine were also obtained and compared with those of the compressor with the original blade angles and the same turbine.

Medeiros, Arthur A.↗
NASA NTRS1954

Analysis of Rocket, Ram-Jet, and Turbojet Engines for Supersonic Propulsion of Long-Range Missles. II - Rocket Missile Performance

The theoretical performance of a two-stage ballistic rocket mis having a centerbody and two parallel boosters was investigated for J oxygen and ammonia-fluorine propellants. Both power-plant and missi parameters were optimized to give minimum cost on-the basis of the analysis for a range of 5500 nautical miles. After optimum values were found, each parameter was varied independently to determine its effect on performance of the missile. The missile using the ammonia-fluorine propellant weighs about one half as much as a missile using JP4-oxygen. Based on an expected unit cost of fluorine in quantity production, the ammonia-fluorine missile has a substantially lower relative cost than a JP4-oxygen missile. Optimum chamber pressures for both propellant systems and for both the centerbody and boosters were between 450 and 600 pounds per square inch. High design altitudes for the exhaust nozzle are desirable for both the centerbody and boosters. For the centerbody, the design altitude should be between 45,000 and 60,000 feet, with the value for ammonia-fluorine lower than that for JP4-oxygen. For the boosters, the design altitude should be 20,000 to 30,000 feet, with the value for the ammonia-fluorine. missile higher.

Huff, Vearl N.↗
NASA NTRS2026

Dynamic Regulation of Sub-Atmospheric Pressure for Constant and Cyclic Gas Loads During Testing of Spacesuit Components

New iterations of the various subsystems within the spacesuit will benefit from a new sorbent technology. For example, sorbents used in the Trace Contaminant Control (TCC) and the Rapid Cycle Amine (RCA) systems within the Exploration Portable Life Support System (xPLSS), part of the Extra-Vehicular Mobility Unit (xEMU). For proper validation, the sub-atmospheric pressure needs to be maintained within a simulated 2 ft 3 spacesuit volume. The traditional approach, which involves placing the entire vent loop within a hypobaric chamber, is not practical for the widespread testing of components and prototypes, as these specialized chambers are costly and not widely available. A two-stage regulator based sub-atmospheric pressure system is presented in this work. The method regulates the pressure based on pressure differentials between the pressure regulation pump, the test system, and the ambient environment. The performance of this system is demonstrated using data collected during a 150+ hour non-regenerative TCC sorbent evaluation, several 24-hour regenerative TCC sorbent evaluations with different regeneration cycles, and an 8-hour RCA sorbent evaluation, all under xEMU operating conditions. For regulation, the inlet regulator was set to propagate a small leak to increase the stability of the system as the gas load from the testing system changes. It was also found that controlling the input flow rate that replaces the material lost during regeneration is critical for ensuring the stability of the system. This system provided excellent pressure regulation, without adjustments, for constant loads during shorter time evaluations (hours), while longer time evaluations (days) are easily obtainable with periodic regulator adjustments. A second method is under development to address longer-term stability and to automate pressure regulation. It will incorporate flow and pressure measurements to dynamically control a set of electronically controlled proportional valves for adjusting the pumping speed and air bleed supply.

Nicholas F Materer↗
DOE OSTI2025

SNoGloDe: A Structured Nonlinear Global Decomposition Solver

Large-scale optimization problems often require decomposition strategies and customized algorithms to achieve optimal solutions within a reasonable time. Building on the work of Cao and Zavala (2019) for solving nonlinear two-stage stochastic programs to global optimality, we implement and extend their approach. We generalize to optimization problems reformulated with a block-angular constraint structure (e.g., temporal decomposition). Our framework, written in Python using Pyomo, is highly customizable and enables parallel execution of the decomposition. SNoGloDe allows tailored branching strategies, lower bounding problems, and candidate generators to leverage problem-specific knowledge. To demonstrate effectiveness, we compare SNoGloDe’s performance with Gurobi on a temporally decomposed produced water case study.

algorithms↗
DOE OSTI2025

A DATA EFFICIENT SPARSE MODELING FRAMEWORK FOR POWER ESTIMATION IN WATER TREATMENT SENSING OPERATIONS

With increasing freshwater scarcity, advanced process design mechanisms such as Closed-Circuit Reverse Osmosis (CCRO) and Digital/Physical Twin systems are gaining traction in water treatment and reuse operations. While digital and physical twin models enable improved system insight and control, their development is often expensive and computationally intensive, requiring large volumes of synthetic or experimental data to characterize underlying process dynamics. This work introduces a sparse surrogate modeling framework to estimate power consumption from measured flow and pressure variables, along with their nonlinear polynomial and interaction expansions. To ensure model reliability and reduce overfitting, a two-stage pipeline is proposed. First, a dynamic data filtering algorithm is employed to remove uninformative observations and transient operational states. Second, a sparse penalized regression technique is applied to select a minimal set of parsimonious features. The proposed model achieves high sparsity, retaining only 7 out of 34 candidate features (≈79.41% sparsity) while delivering a root mean square error (RMSE) of 0.072 on the test dataset.

Mukherjee, Subrata [ORNL] (ORCID:0000000309930338)↗
DOE OSTI2026

Estimating Flexibility Envelopes for Residential Customers From Utility Smart Meter Data: Preprint

Demand response from residential customers has significant potential to support power system operations, but accurate flexibility estimation is challenging due to the limited resolution of advanced metering infrastructure (AMI) data. Most utility AMI measurements are recorded at hourly intervals, with only a small portion at higher resolutions, and even fewer households have appliance-level energy usage data. To address this issue, this paper proposes a two-stage long short-term memory (LSTM) framework for estimating household flexibility envelopes from low-resolution AMI data. In the first stage, the heating, ventilating, and air-conditioning (HVAC) load and non-HVAC loads are estimated by using a model trained on a small set of households with appliance-level profiles. These estimated data are then used to compute the upper- and lower-flexibility bounds, which are subsequently down-sampled to lower-resolution data. In the second stage, these flexibility bounds serve as training inputs for another LSTM model, enabling direct prediction of flexibility envelopes for households with only hourly AMI data. This method is validated using Pecan Street data from two different areas, and the results demonstrate its applicability and effectiveness.

24 POWER TRANSMISSION AND DISTRIBUTION↗