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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 1,009 records · Page 56
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Transient Three-Dimensional Side Load Analysis of Out-of-Round Film Cooled Nozzles

The objective of this study is to investigate the effect of nozzle out-of-roundness on the transient startup side loads at a high altitude, with an anchored computational methodology. The out-of-roundness could be the result of asymmetric loads induced by hardware attached to the nozzle, asymmetric internal stresses induced by previous tests, and deformation, such as creep, from previous tests. The rocket engine studied encompasses a regeneratively cooled thrust chamber and a film cooled nozzle extension with film coolant distributed from a turbine exhaust manifold. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Transient startup computations were performed with the out-of-roundness achieved by four different degrees of ovalization: one perfectly round, one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation-line-jump is the peak side load physics for the round, slightly our-of-round, and more out-of-round cases, and the peak side load increases as the degree of out-of-roundness increases. For the significantly out-of-round nozzle, however, the peak side load reduces to comparable to that of the round nozzle and the separation line jump is not the peak side load physics. The counter-intuitive result of the significantly out-of-round case is found to be related to a side force reduction mechanism that splits the effect of the separation-line-jump into two parts, not only in the circumferential direction and most importantly in time.

Wang, Ten-See↗
NASA NTRS2011

Flight Test Results of a GPS-Based Pitot-Static Calibration Method Using Output-Error Optimization for a Light Twin-Engine Airplane

As part of the NASA Aviation Safety Program (AvSP), a novel pitot-static calibration method was developed to allow rapid in-flight calibration for subscale aircraft while flying within confined test areas. This approach uses Global Positioning System (GPS) technology coupled with modern system identification methods that rapidly computes optimal pressure error models over a range of airspeed with defined confidence bounds. This method has been demonstrated in subscale flight tests and has shown small 2- error bounds with significant reduction in test time compared to other methods. The current research was motivated by the desire to further evaluate and develop this method for full-scale aircraft. A goal of this research was to develop an accurate calibration method that enables reductions in test equipment and flight time, thus reducing costs. The approach involved analysis of data acquisition requirements, development of efficient flight patterns, and analysis of pressure error models based on system identification methods. Flight tests were conducted at The University of Tennessee Space Institute (UTSI) utilizing an instrumented Piper Navajo research aircraft. In addition, the UTSI engineering flight simulator was used to investigate test maneuver requirements and handling qualities issues associated with this technique. This paper provides a summary of piloted simulation and flight test results that illustrates the performance and capabilities of the NASA calibration method. Discussion of maneuver requirements and data analysis methods is included as well as recommendations for piloting technique.

Martos, Borja↗
NASA NTRS2008

Jet Engine Fan Blade Containment Using an Alternate Geometry

With a goal of reducing jet engine weight, simulations of a fan blade containment system with an alternate geometry were tested and analyzed. A projectile simulating a fan blade was shot at two alternate geometry containment case configurations using a gas gun. The first configuration was a flat plate representing a standard case configuration. The second configuration was a flat plate with a radially convex curve section at the impact point. The curved surface was designed to force the blade to deform plastically, dissipating energy before the full impact of the blade is received by the plate. The curved case was able to tolerate a higher impact velocity before failure. The computational model was developed and correlated with the tests and a weight savings assessment was performed. For the particular test configuration used in this study the ballistic impact velocity of the curved plate was approximately 60 m/s (200 ft/s) greater than that of the flat plate. For the computational model to successfully duplicate the test, the very high strain rate behavior of the materials had to be incorporated.

Carney, K.S.↗
NASA NTRS2012

Transient Side Load Analysis of Out-of-Round Film-Cooled Nozzle Extensions

There was interest in understanding the impact of out-of-round nozzle extension on the nozzle side load during transient startup operations. The out-of-round nozzle extension could be the result of asymmetric internal stresses, deformation induced by previous tests, and asymmetric loads induced by hardware attached to the nozzle. The objective of this study was therefore to computationally investigate the effect of out-of-round nozzle extension on the nozzle side loads during an engine startup transient. The rocket engine studied encompasses a regeneratively cooled chamber and nozzle, along with a film cooled nozzle extension. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and transient inlet boundary flow properties derived from an engine system simulation. Six three-dimensional cases were performed with the out-of-roundness achieved by three different degrees of ovalization, elongated on lateral y and z axes: one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation line jump was the primary source of the peak side loads. Comparing to the peak side load of the perfectly round nozzle, the peak side loads increased for the slightly and more ovalized nozzle extensions, and either increased or decreased for the two significantly ovalized nozzle extensions. A theory based on the counteraction of the flow destabilizing effect of an exacerbated asymmetrical flow caused by a lower degree of ovalization, and the flow stabilizing effect of a more symmetrical flow, created also by ovalization, is presented to explain the observations obtained in this effort.

Wang, Ten-See↗
NASA NTRS2015

ME8373 Spring 2015 ICME Proposal ICME Analysis of Fatigue Crack Growth Through a Weld in SA-516 Grade 70 Plate

This paper describes work accomplished to predict the service life of a flexure joint design which is a component of a diffuser duct in the A3 Test Stand, an altitude simulation rocket engine test facility at NASA's Stennis Space Center. The duct has two pressure shells separated by cooling water passages and connected by stiffening ribs and flexure joints. Rocket exhaust flows within the duct and heats the inner pressure shell while the outer pressure shell remains at ambient temperature. The flexure joints allow for differential thermal expansion of the inner and outer pressure shells and are subject to in-service loading by this thermal expansion along with water pressure in the cooling water passage, atmospheric pressure outside the duct, near vacuum conditions within the duct, and vibrational loads from operation of the facility and rocket engine. Figure 1 shows a schematic axisymmetric cross section of the diffuser pressure shells and flexure joints with a zoomed in view of the flexure joint. The flexure joints are expected to eventually fail by fatigue cracking leading to leaks from the cooling water passages to the outside. The zoomed in view in Figure 1 indicates where cracking is expected to occur, namely through a weld bead between two plates of SA-516 Grade 70 steel. This weld bead acts as the fulcrum of the flexure joint and it is clear from inspection of the geometry and loading represented in the zoomed in portion of Figure 1 that inherent in the design there is a severe notch formed between the flexure plate, weld bead, and stiffening ring that will be the site of crack initiation and location from which the crack grows to the outer surface of the weld bead.

Woods, Jody L.↗
NASA NTRS2016

Alternative Bio-Derived JP-8 Class Fuel and JP-8 Fuel: Flame Tube Combustor Test Results Compared using a GE TAPS Injector Configuration

This paper presents results from tests in a flame tube facility, where a bio-derived alternate fuel was compared with JP-8 for emissions and general combustion performance. A research version of General Electric Aviation (GE) TAPS injector was used for the tests. Results include combustion efficiency from gaseous emission measurements, 2D planar laser-based imaging as well as basic flow visualization of the flame. Four inlet test conditions were selected that simulate various engine power conditions relevant to NASA Fundamental Aeronautics Supersonics Project and Environmentally Responsible Aviation Program. One inlet condition was a pilot-only test point. The other three inlet conditions incorporated fuel staging via a split between the pilot and main circuits of either 10%/90% or 20%/80%. For each engine power condition, three fuel mixes were used: 100% JP-8; 100% alternative; and a blend of the two, containing 75% alternative. Results for the inlet cases that have fuel split between pilot and main, indicate that fuel from the pilot appears to be evaporated by the time it reaches the dome exit. Main circuit liquid evaporates within a downstream distance equal to annulus height, no matter the fuel. Some fuel fluorescence images for a 10%/90% fuel staging case show a distinct difference between JP-8 and bio-derived fuel. OH PLIF results indicate that OH forms in a region more centrally-located for the JP-8 case downstream of the pilot, in its central recirculation region (CRZ). For the bio-derived Hydrotreated Renewable Jet (HRJ) fuel, however, we do not see much OH in the CRZ. The OH image structure near the dome exit is similar for the two fuels, but farther downstream the OH in the CRZ is much more apparent for the JP-8 than for the alternate fuel. For all conditions, there was no discernable difference between fuel types in combustion efficiency or emissions.

gas turbine combustors↗
NASA NTRS2016

Alternative Bio-Derived JP-8 Class Fuel and JP-8 Fuel: Flame Tube Combustor Test Results Compared using a GE TAPS Injector Configuration

This paper presents results from tests in a NASA Glenn Research Center (GRC) flame tube facility, where a bio-derived alternate fuel was compared with JP-8 for emissions and general combustion performance. A research version of General Electric Aviation (GE) TAPS injector was used for the tests. Results include 2D, planar laser-based imaging as well as basic flow visualization of the flame. Four conditions were selected that simulate various engine power conditions relevant to NASA Fundamental Aeronautics Supersonics and Environmentally Responsible Aviation Projects were tested.

optical diagnostics↗
NASA NTRSDate not supplied

Bootstrapping Multi-Agent Unmanned Aerial Vehicle (UAV) System Integration Using Ground-Based Assets: Lessons Learned

The highly dynamic nature of UAVs imposes significant challenges when conducting initial testing ranging from safety risks posed by high-capacity lithium batteries and spinning propellers to rigorous timing demands on controllers and the consequences of failures mid-air. Flight testing of a single vehicle is time and labor intensive due to these challenges and more, and the complexity increases exponentially with the number of vehicles. While simulations and hardware-in-the-loop bench testing can provide adequate environments for preliminary validation, differences in system deployment architecture, software interfaces, and hardware infrastructure between simulation and a fleet of real UAVs create a sizable gap that must be navigated carefully during system integration. In support of the Autonomy Teaming and TRAjectories for Complex Trusted Operational Reliability (ATTRACTOR) project, which had the goal of establishing a basis of certification of trust and trustworthiness in multi-agent autonomous systems, this gap was tackled from two directions. First, a novel mixed-reality simulation environment was engineered to blur the transition from simulation to flight hardware. Second, a fleet of Unmanned Surface Vehicles (USVs) was developed as a test and evaluation platform that more closely represented the final aerial fleet while eliminating many of the risks associated with air vehicles. This paper delves into the second element, analyzing the efficacy of the USV platform in performing system integration testing for the UAV system. In this paper we present the USV fleet and its role in reducing the aforementioned gaps in deployment architecture, software interfaces, and hardware infrastructure when moving from simulation to flight. An overview of the hardware and software onboard the vehicles will be provided along with supporting infrastructure. The system integration process will be documented including results in supporting both the overarching design reference mission (DRM) of ATTRACTOR and individual research efforts conducted during the project. Finally, we will discuss some of the practical lessons learned regarding the testing, deployment, and operation of multi-agent autonomous systems.

Matthew P Vaughan↗
NASA NTRSDate not supplied

Hybrid Propulsion Emulation Rig (HyPER)

HyPER is a hardware-in-the-loop laboratory that was designed specifically to investigate the dynamic interactions between turbomachinery, the electric power system, and the constantly varying loads of electrified aircraft. It is a small-scale lab capable of rapid reconfiguration through software. This allows the emulation of new engines using simulation models that are easily replaced and then appropriately scaled for power and inertia to the test hardware.

Control↗
NASA NTRS2020

Experimental Validation of Nozzle Flow Simulations for Rotating Detonation Rocket Engines

Rotating detonation engines (RDEs) promise increased thermodynamic performance that may significantly enhance the capabilities of current rocket platforms. Little work has been conducted thus far to characterize the effect of nozzle design on high chamber pressure RDEs. Previous computational work was completed to understand the nozzle performance of the Purdue methane/oxygen rocket RDE. A new experimental study on a similar kerosene/oxygen RDE was conducted to validate the results of the computational study. Both a nozzleless geometry and several aerospike designs were hot-fire tested. New pressure instrumentation on the different nozzle surfaces were included to better understand the flow physics unique to RDE chamber exit conditions. Both the previous computational study and the new experimental results confirmed that for a nozzleless geometry, the RDE cycle enhances suction on the base region, an important result to determining RDE engine performance separate from nozzle effects. While the aerospike experiments showed agreement with computational results, delay of flow separation due to the RDE cycle could not be confirmed. Two aerospike geometries with different nozzle pressure ratios were experimentally evaluated. This study showed both the strength and the necessity for 3D transient computations to better understand the RDE flow field with nozzle geometries.

Alexis J Harroun↗
NASA NTRSDate not supplied

Contrails Measurement and Testing Capabilities in NASA's Particulate Aerosol Laboratory

The NASA Glenn Research Center Particulate Aerosol Laboratory (PAL) is a ground test facility for studying emissions and contrails at upper atmospheric conditions. The facility consists of a small-scale combustor that generates hot combustion gases and soot nanoparticles, which flow through a transition pipe into a jet nozzle, which exhausts into an altitude chamber. Thermodynamic conditions experienced in flight are matched in the altitude chamber, allowing the study of ice particle formation of simulated aircraft engine exhaust plumes. A non-intrusive optical diagnostic technique based on Mie scattering provides ice particle size, number density, and mass density. A suite of commercial extractive particle measurement instruments provides estimates of soot number density, size, and black carbon mass distributions. Measurements are presented for various fuels at chamber conditions covering an altitude range of 20,000 ft (6 km) to 45,000 ft (14 km) altitude and ambient temperatures from -48°C to -68°C. The data presented demonstrate the effect of soot number density, altitude, and fuel-to-air equivalence ratio on the contrail ice particle formation. Ice mass density measurements are compared to values predicted by a thermodynamic plume similarity solution.

Optical Instrumentation↗
NASA NTRS2026

Optimization of a 10–Step Reduced CH₄/O₂ Combustion Mechanism for RDRE Applications

A 10-step, 9-species reduced mechanism for methane/oxygen combustion is evaluated and optimized for use in rotating detonation rocket engine (RDRE) simulations at an operating pressure of 10 bar. The mechanism is assessed against GRI-Mech 3.0 for ignition delay times, adiabatic equilibrium temperature and species, Chapman-Jouguet (CJ) detonation parameters, and laminar flame speed. A sequential Nelder–Mead optimization procedure is applied to the rate constants of reactions R1, R4, and R5, targeting ignition delay agreement over the equivalence ratio range φ = 1.0–1.4 and temperature range T = 1400–2000 K. The optimized mechanism achieves ignition delay agreement within 28% of GRI-Mech 3.0 for the temperature range relevant to RDRE operation (τ ≥ 5 μs). Adiabatic equilibrium temperatures and CJ detonation parameters are reproduced within 0.15% of GRI-Mech 3.0 values, confirming thermodynamic consistency of the mechanism. Laminar flame speed predictions deviate significantly from GRI-Mech 3.0, which is attributed to the absence of HO₂ chemistry in the reduced mechanism; this limitation does not affect RDRE simulation fidelity since laminar flame propagation is of secondary importance in RDREs.

Computational Fluid Dynamics↗
NASA NTRS2010

Rocket-Plume Spectroscopy Simulation for Hydrocarbon-Fueled Rocket Engines

The UV-Vis spectroscopic system for plume diagnostics monitors rocket engine health by using several analytical tools developed at Stennis Space Center (SSC), including the rocket plume spectroscopy simulation code (RPSSC), to identify and quantify the alloys from the metallic elements observed in engine plumes. Because the hydrocarbon-fueled rocket engine is likely to contain C2, CO, CH, CN, and NO in addition to OH and H2O, the relevant electronic bands of these molecules in the spectral range of 300 to 850 nm in the RPSSC have been included. SSC incorporated several enhancements and modifications to the original line-by-line spectral simulation computer program implemented for plume spectral data analysis and quantification in 1994. These changes made the program applicable to the Space Shuttle Main Engine (SSME) and the Diagnostic Testbed Facility Thruster (DTFT) exhaust plume spectral data. Modifications included updating the molecular and spectral parameters for OH, adding spectral parameter input files optimized for the 10 elements of interest in the spectral range from 320 to 430 nm and linking the output to graphing and analysis packages. Additionally, the ability to handle the non-uniform wavelength interval at which the spectral computations are made was added. This allowed a precise superposition of wavelengths at which the spectral measurements have been made with the wavelengths at which the spectral computations are done by using the line-by-line (LBL) code. To account for hydrocarbon combustion products in the plume, which might interfere with detection and quantification of metallic elements in the spectral region of 300 to 850 nm, the spectroscopic code has been enhanced to include the carbon-based combustion species of C2, CO, and CH. In addition, CN and NO have spectral bands in 300 to 850 nm and, while these molecules are not direct products of hydrocarbon-oxygen combustion systems, they can show up if nitrogen or a nitrogen compound is present as an impurity in the propellants and/or these can form in the boundary layer as a result of interaction of the hot plume with the atmosphere during the ground testing of engines. Ten additional electronic band systems of these five molecules have been included into the code. A comprehensive literature search was conducted to obtain the most accurate values for the molecular and the spectral parameters, including Franck-Cordon factors and electronic transition moments for all ten band systems. For each elemental transition in the RPSSC, six spectral parameters - Doppler broadened line width at half-height, pressure-broadened line width at half-height, electronic multiplicity of the upper state, electronic term energy of the upper state, Einstein transition probability coefficient, and the atomic line center - are required. Input files have been created for ten elements of Ni, Fe, Cr, Co, Cu, Ca, Mn, Al, Ag, and Pd, which retain only relatively moderate to strong transitions in 300 to 430 nm spectral range for each element. The number of transitions in the input files is 68 for Ni; 148 for Fe; 6 for Cr; 87 for Co; 1 for Ca; 3 for Mn; 2 each for Cu, Al, and Ag; and 11 for Pd.

Tejwani, Gopal D.↗
NASA NTRS1981

Exhaust emissions survey of a turbofan engine for flame holder swirl type augmentors at simulated altitude flight conditions

Emissions of carbon dioxide, total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from an F100 afterburning two spool turbofan engine at simulated flight conditions are reported. Tests were run at Mach 0.8 at altitudes of 10.97 and 13.71 km (36,000 and 45,000 ft), and at Mach 1.2 at 13.71 km (45,000 ft). Emission measurements were made from intermediate power (nonafterburning) through maximum afterburning, using a single point gas sample probe traversed across the horizontal diameter of the exhaust nozzle. The data show that emissions vary with flight speed, altitude, power level, and radial position across the nozzle. Carbon monoxide emissions were low for intermediate and partial afterburning power. Unburned hydrocarbons were near zero for most of the simulated flight conditions. At maximum afterburning, there were regions of NOx deficiency in regions of high CO. The results suggest that the low NOx levels observed in the tests are a result of interaction with high CO in the thermal converter. CO2 emissions were proportional to local fuel air ratio for all test conditions.

Moss, J. E., Jr.↗
NASA NTRS2026

NASA JSC’s Simulant Development Lab Capabilities and Artemis Testing

The Simulant Development Lab (SDL) is a multifunctional collaborative workspace that supports the development, curation, analysis, testing, and distribution of planetary regolith simulants – including lunar, Martian, asteroidal, and other granular materials. The lab provides a multidisciplinary setting for scientific characterization of simulant physical properties and for engineering evaluations conducted with simulant test beds. To enable this work, the SDL curates and maintains a stock of more than 35 metric tons of simulant material. To evaluate these materials and support testing goals, the lab is equipped with a comprehensive suite of processing tools and analytical instruments. These capabilities enable the SDL’s mission at NASA’s Johnson Space Center to distribute, develop, process, characterize, and test regolith simulants for mission relevant applications. Through controlled and repeatable testing environments that replicate the physical and compositional properties of lunar regolith, the SDL supports Artemis hardware maturation, providing safe, Earth‑based analogs for evaluating systems that must withstand regolith dust interactions, physical wear and abrasion, and operational loads. The facility’s extensive simulant inventory and integrated geological and engineering test infrastructure accelerate technology readiness for Artemis and future exploration campaigns (e.g., future crewed or robotic missions to Mars).

Simulant Development Lab↗
NASA NTRS1985

Man-vehicle systems research facility advanced aircraft flight simulator throttle mechanism

The Advanced Aircraft Flight Simulator is equipped with a motorized mechanism that simulates a two engine throttle control system that can be operated via a computer driven performance management system or manually by the pilots. The throttle control system incorporates features to simulate normal engine operations and thrust reverse and vary the force feel to meet a variety of research needs. While additional testing to integrate the work required is principally now in software design, since the mechanical aspects function correctly. The mechanism is an important part of the flight control system and provides the capability to conduct human factors research of flight crews with advanced aircraft systems under various flight conditions such as go arounds, coupled instrument flight rule approaches, normal and ground operations and emergencies that would or would not normally be experienced in actual flight.

Kurasaki, S. S.↗
NASA NTRS1984

Computer simulation of the heavy-duty turbo-compounded diesel cycle for studies of engine efficiency and performance

Reductions in heat loss at appropriate points in the diesel engine which result in substantially increased exhaust enthalpy were shown. The concepts for this increased enthalpy are the turbocharged, turbocompounded diesel engine cycle. A computer simulation of the heavy duty turbocharged turbo-compounded diesel engine system was undertaken. This allows the definition of the tradeoffs which are associated with the introduction of ceramic materials in various parts of the total engine system, and the study of system optimization. The basic assumptions and the mathematical relationships used in the simulation of the model engine are described.

Assanis, D. N.↗