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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Calibration for Thrust and Airflow Measurements in the CE-22 Advanced Nozzle Test Facility

CE-22 facility procedures and measurements for thrust and airflow calibration obtained with choked-flow ASME nozzles are presented. Six calibration nozzles are used at an inlet total pressure from 20 to 48 psia. Throat areas are from 9.9986 to 39.986 sq. in.. Throat Reynolds number varies from 1.8 to 7.9 million. Nozzle gross thrust coefficient (CFG) uncertainty is 0.25 to 0.75 percent, with smaller uncertainly generally for larger nozzles and higher inlet total pressure. Nozzle discharge coefficient (CDN) uncertainty is 0.15 percent or less for all the data. ASME nozzle calibrations need to be done before and after research model testing to achieve these uncertainties. In addition, facility capability in terms of nozzle pressure ratio (NPR) and nozzle airflow are determined. Nozzle pressure ratio of 50 or more is obtainable at 40 psia for throat areas between 20 and 30 sq. in.. Also presented are results for two of the ASME nozzles vectored at 10deg, a dead-weight check of the vertical (perpendicular to the jet axis) force measurement, a calibration of load cell forces for the effects of facility tank deflection with tank pressure, and the calibration of the metric-break labyrinth seal.

Werner, Roger A.↗
NASA NTRS1983

Effects of wind on turbofan engines in outdoor static test stands

Wind can affect measured thrust and can cause turbofan engine speed to fluctuate during outdoor testing. Techniques used at an outdoor test stand at NASA Lewis Research Center to make testing easier and faster and to improve data repeatability include using an inflow control device (ICD) to make fan speed steadier, taking many raw data samples for better averaging, and correcting thrust for wind direction and speed. Data from engine tests are presented to show that the techniques improve repeatability of thrust and airflow measurements under various wind conditions.

Mcardle, J. G.↗
NASA NTRS1983

Effects of wind on turbofan engines in outdoor static test stands

Wind can affect measured thrust and can cause turbofan engine speed to fluctuate during outdoor testing. Techniques used at an outdoor test stand at NASA Lewis Research Center to make testing easier and faster and to improve data repeatability include using an inflow control device (ICD) to make fan speed steadier, taking many raw data samples for better averaging, and correcting thrust for wind direction and speed. Data from engine tests are presented to show that the techniques improve repeatability of thrust and airflow measurements under various wind conditions. Previously announced in STAR as N83-34945

Mcardle, J. G.↗
NASA NTRS2005

Effect of Operating Frequency and Fill Time on PDE-Ejector Thrust Performance

Thrust measurements for a pulse detonation engine (PDE)-ejector system were determined for a range of operating frequencies. Various length tubular ejectors were utilized. The results were compared to the measurements of the thrust output of the PDE alone to determine the enhancement provided by each ejector configuration at the specified frequencies. Ethylene was chosen as the fuel, with an equi-molar mixture of nitrogen and oxygen acting as the oxidizer. The propellant was kept at an equivalence ratio of one during all the experiments. The system was operated for frequencies between 20 and 50 Hz. The parameter space of the study included PDE operation frequency, ejector length, overlap percentage, the radius of curvature for the ejector inlets, and duration of the time allowed between cycles. The results of the experiments showed a maximum thrust augmentation of 120% for a PDE-ejector configuration at a frequency of 40Hz with a fill time of 10 ms.

Landry, K.↗
NASA NTRS1989

Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices was designed and tested. The thrust stand was specifically tailored to the needs of a 100 to 250 kW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. A detailed description is given for the thrust stand design and operation with a 100 kW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability.

Haag, Thomas W.↗
NASA NTRS2005

Studies of Operating Frequency Effects On Ejector-based Thrust Augmentation in a Pulse Detonation Engine

Studies were performed in order to characterize the thrust augmentation potential of an ejector in a Pulse Detonation Engine application. A 49-mm diameter tube of 0.914-m length was constructed with one open end and one closed end. Ethylene, oxygen, and nitrogen were introduced into the tube at the closed end through the implementation of a fast mixing injector. The tube was completely filled with a stoichiometric mixture containing a one to one molar ratio of nitrogen to oxygen. Ethylene was selected as the fuel due to its detonation sensitivity and the molar ratio of the oxidizer was chosen for heat transfer purposes. Detonations were initiated in the tube through the use of a spark ignition system. The PDE was operated in a multi-cycle mode at frequencies ranging from 20-Hz to 50-Hz. Baseline thrust measurements with no ejector present were performed while operating the engine at various frequencies and compared to theoretical estimates. The baseline values were observed to agree with the theoretical model at low operating frequencies and proved to be increasingly lower than the predicted values as the operating frequency was increased. The baseline thrust measurements were observed to agree within 15 percent of the model for all operating frequencies. A straight 152-mm diameter ejector was installed and thrust augmentation percentages were measured. The length of the ejector was varied while the overlap percentage (percent of the ejector length which overlapped the tube) was maintained at 25 percent for all tests. In addition, the effect of ejector inlet geometry was investigated by comparing results with a straight inlet to those of a 38-mm inlet diameter. The thrust augmentation of the straight inlet ejector proved to be independent of engine operating frequency, augmenting thrust by 40 percent for the 0.914-m length ejector. In contrast, the rounded lip ejector of the same length seemed to be highly dependent on the engine operating frequency. An optimum operating frequency observed with the rounded inlet occurred at an operating frequency of 30-Hz, resulting in thrust augmentation percentages greater than 100 percent. The effect that the engine operating frequency had on thrust augmentation levels attained with an ejector was characterized and optimum performance parameters were established. Insight into the frequency dependent nature of the ejector performance was pursued. Suggestions for future experiments which are needed to fully understand the means in which thrust augmentation is achieved in a PDE-ejector configuration were noted.

Landry, K.↗
NASA NTRS1989

Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices has been designed and tested. The thrust stand was specifically tailored to the needs of a 0.1 to 0.25 MW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. The paper gives a detailed description of the thrust stand design and operation with a 0.1 MW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability.

Haag, Thomas W.↗
NASA NTRS2005

Experimental Study of a Pulse Detonation Engine Driven Ejector

Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen (O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results are in excellent agreement with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. Further increases are possible by tailoring the ejector geometry based on CFD predictions conducted elsewhere. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Santoro, Robert J.↗
NASA NTRS2004

Development and Demonstration of a Device to Determine Thrust by Measuring the Force on a Target Plate in the Exhaust of a Plasma Thruster

A device has been developed to measure the force on a target plate by an impacting beam of charged and neutral particles. This device, an impact thrust stand, was developed to allow thrusters at low TRL, levels to be easily tested without the expense of developing a flight prototype of the thruster to be placed on a conventional thrust stand. The impact thrust stand was developed for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) but has been tested and calibrated using several devices including Hall thrusters. The calibration and comparison of the impact thrust stand against conventional thrust stands will be discussed in this paper.

Chavers, Greg↗
NASA NTRS1988

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1030:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

Smith, Tamara A.↗
NASA NTRS1988

Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1039:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

Smith, Tamara A.↗
NASA NTRS1971

Experimental techniques for evaluating steady-state jet engine performance in an altitude facility

Jet engine calibration tests were conducted in an altitude facility using a contoured bellmouth inlet duct, four fixed-area water-cooled exhaust nozzles, and an accurately calibrated thrust measuring system. Accurate determination of the airflow measuring station flow coefficient, the flow and thrust coefficients of the exhaust nozzles, and the experimental and theoretical terms in the nozzle gross thrust equation were some of the objectives of the tests. A primary objective was to develop a technique to determine gross thrust for the turbojet engine used in this test that could also be used for future engine and nozzle evaluation tests. The probable error in airflow measurement was found to be approximately 0.6 percent at the bellmouth throat design Mach number of 0.6. The probable error in nozzle gross thrust measurement was approximated 0.6 percent at the load cell full-scale reading.

Smith, J. M.↗
NASA NTRS1985

An analysis of cross-coupling of a multicomponent jet engine test stand using finite element modeling techniques

A two axis thrust measuring system was analyzed by using a finite a element computer program to determine the sensitivities of the thrust vectoring nozzle system to misalignment of the load cells and applied loads, and the stiffness of the structural members. Three models were evaluated: (1) the basic measuring element and its internal calibration load cells; (2) the basic measuring element and its external load calibration equipment; and (3) the basic measuring element, external calibration load frame and the altitude facility support structure. Alignment of calibration loads was the greatest source of error for multiaxis thrust measuring systems. Uniform increases or decreases in stiffness of the members, which might be caused by the selection of the materials, have little effect on the accuracy of the measurements. It is found that the POLO-FINITE program is a viable tool for designing and analyzing multiaxis thrust measurement systems. The response of the test stand to step inputs that might be encountered with thrust vectoring tests was determined. The dynamic analysis show a potential problem for measuring the dynamic response characteristics of thrust vectoring systems because of the inherently light damping of the test stand.

Schweikhard, W. G.↗
NASA NTRS2013

Inductive Pulsed Plasma Thruster Development and Testing at NASA-MSFC

THE inductive pulsed plasma thruster (IPPT) is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. In the present work, we present a summary of the IPPT research and development conducted at NASA's Marshall Space Flight Center (MSFC). As a higher-power, still relatively low readiness level system, there are many issues associated with the eventual deployment and use of the IPPT as a primary propulsion system on spacecraft that remain to be addressed. The present program aimed to fabricate and test hardware to explore how these issues could be addressed. The following specific areas were addressed within the program and will be discussed within this paper. a) Conical theta-pinch IPPT geometry thruster configuration. b) Repetition-rate multi-kW thruster pulsing. c) Long-lifetime pulsed gas valve. d) Fast pulsed gas valve driver and controller. e) High-voltage, repetitive capacitor charging power processing unit. During the course of testing, a number of specific tests were conducted, including several that, to our knowledge, have either never been previously conducted (such as multi-KW repetition-rate operation) or have not been performed since the early 1990s (direct IPPT thrust measurements).2 Conical theta-pinch IPPT thrust stand measurements are presented in Fig. 1 while various time-integrated and time

Polzin, Kurt A.↗
NASA NTRS2015

Electronegative Gas Thruster

The project is an international collaboration and academic partnership to mature an innovative electric propulsion thruster concept to Technology Research Level-3 (TRL-3) through direct thrust measurement. The project includes application assessment of the technology ranging from small spacecraft to high power. The Plasma propulsion with Electronegative GASES(PEGASES) basic proof of concept has been matured to TRL-2 by Ane Aanesland of Laboratoire de Physique des Plasma at Ecole Polytechnique. The concept has advantages through eliminating the neutralizer requirement and should yield longer life and lower cost over conventional gridded ion engines. The objective of this research is to validate the proof of concept through the first direct thrust measurements and mature the concept to TRL-3.

Dankanich, John↗
NASA NTRS2010

Thrust Stand Measurements of the Conical Theta Pinch FARAD Thruster

It is found that the impulse of a pulsed inductive plasma thruster utilizing preionization is maximized for a particular ratio of the stored energy in the capacitor to the injected propellant mass. The fact that the impulse depends on the ratio of the initial stored energy to injected propellant mass agrees with previous current sheet studies, supporting the idea that a Townsend-like breakdown process strongly influences current sheet formation, and in turn, current sheet formation strongly affects the operational efficiency of the device. The optimum in half cone angle of the inductive coil can be explained in terms of a balance between the direct axial acceleration and the radial pinching contribution to thrust. From the trends in these data we conclude that operation at the correct ratio of capacitor energy to propellant mass is essential for efficient operation of pulsed inductive plasma thrusters employing a preionized propellant.

Hallock, Ashley K.↗