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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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

QFD analysis of RSRM aqueous cleaners

This paper presents a Quality Function Deployment (QFD) analysis of the final down-selected aqueous cleaners to be used on the Redesigned Solid Rocket Motor (RSRM) program. The new cleaner will replace solvent vapor degreasing. The RSRM Ozone Depleting Compound Elimination program is discontinuing the methyl chloroform vapor degreasing process and replacing it with a spray-in-air aqueous cleaning process. Previously, 15 cleaners were down-selected to two candidates by passing screening tests involving toxicity, flammability, cleaning efficiency, contaminant solubility, corrosion potential, cost, and bond strength. The two down-selected cleaners were further evaluated with more intensive testing and evaluated using QFD techniques to assess suitability for cleaning RSRM case and nozzle surfaces in preparation for adhesive bonding.

Marrs, Roy D.↗
NASA NTRS2010

Supersonic Free-Jet Combustion in a Ramjet Burner

A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is governed largely by the same physics as its classical counterpart. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode configurations. Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD) analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and 12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions between the free-jet and recirculating combustion chamber flows. Performance based on these CFD results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these interactions. These differences are quantified and discussed. Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation on combustion chamber pressure, flow structure, and performance. Calculations with constant temperature walls were also done to evaluate heat flux and overall heat loads. Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and combustion chamber pressurization. Also recommended is an examination of system-level aspects such as weight, thermal management and rocket integration as well as alternate geometries and variable geometry schemes.

Trefny, Charles J.↗
NASA NTRS1981

Technology development status at McDonnell Douglas

The significant technology items of the Concorde and the conceptual MCD baseline advanced supersonic transport are compared. The four major improvements are in the areas of range performance, structures (materials), aerodynamics, and in community noise. Presentation charts show aerodynamic efficiency; the reoptimized wing; low scale lift/drag ratio; control systems; structural modeling and analysis; weight and cost comparisons for superplasticity diffusion bonded titanium sandwich structures and for aluminum brazed titanium honeycomb structures; operating cost reduction; suppressor nozzles; noise reduction and range; the bicone inlet; a market summary; environmental issues; high priority items; the titanium wing and fuselage test components; and technology validation.

Rowe, W. T.↗
NASA NTRS1993

An experimental investigation of supersonic swirling jets

The addition of swirl to scramjet fuel jets has been proposed as a method of enhancing fuel mixing, but little of a fundamental nature is known about supersonic swirling flows. This is a study of mixing in supersonic swirling air jets created by tangential injection and acceleration through a convergent-divergent nozzle into stagnant air. Several cases with different vortex circulations were investigated using pitot, cone and total-temperature probes, and Rayleigh scattering from a laser light sheet. The results show that tangential injection is an efficient method for generating swirling jets; that the swirling jets mix much more rapidly with the ambient air than comparable straight jets; and that when overexpanded, considerable turbulence and increased mixing is created as a result of vortex breakdown.

Cutler, Andrew D.↗
NASA NTRS1951

Combustion-chamber Performance Characteristics of a Python Turbine-propeller Engine Investigated in Altitude Wind Tunnel

Combustion-chamber performance characteristics of a Python turbine-propeller engine were determined from investigation of a complete engine over a range of engine speeds and shaft horsepowers at simulated altitudes. Results indicated the effect of engine operating conditions and altitude on combustion efficiency and combustion-chamber total pressure losses. Performance of this vaporizing type combustion chamber was also compared with several atomizing type combustion chambers. Over the range of test conditions investigated, combustion efficiency varied from approximately 0.95 to 0.99.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1952

Experimental Investigation of Flow in an Annular Cascade of Turbine Nozzle Blades of Constant Discharge Angle

The experimental performance of turbine nozzle blades designed for a constant discharge angle was investigated at discharge hub Mach numbers of 1.18, 1.31, and 1.41. Flow characteristics are presented in terms of energy losses, angle gradients, and secondary flow effects. Blade efficiency decreased from 0.983 to 0.978 with increasing Mach number in the range investigated while angle variations in the loss regions became very large, indicating poorer blade performance than efficiency implies.

BOUNDARY LAYER CHARACTERISITCS-INTERNAL AERODYNAMI↗
NASA NTRS1952

Investigation at Mach Numbers 2.98 and 2.18 of Axially Symmetric Free-jet Diffusion with a Ram-jet Engine

An investigation was conducted to determine the effectiveness of a free-jet diffuser in reducing the over-all pressure ratios required to operate a free jet with a large air-breathing engine as a test vehicle. Efficient operation of the free jet was determined with and without the considerations required for producing suitable engine-inlet flow conditions. A minimum operating pressure ration of 5.5 was attained with a ratio of nozzle-exit to engine-inlet area of 1.85. Operation of the free jet with unstable engine-inlet flow (buzz) is also included.

WIND TUNNELS↗
NASA NTRS1999

Aerothermo-Structural Analysis of Low Cost Composite Nozzle/Inlet Components

This research is a cooperative effort among the Turbomachinery and Propulsion Division of NASA Glenn, CCMR of NC A&T State University, and the Tuskegee University. The NC A&T is the lead center and Tuskegee University is the participating institution. Objectives of the research were to develop an integrated aerodynamic, thermal and structural analysis code for design of aircraft engine components, such as, nozzles and inlets made of textile composites; conduct design studies on typical inlets for hypersonic transportation vehicles and setup standards test examples and finally manufacture a scaled down composite inlet. These objectives are accomplished through the following seven tasks: (1) identify the relevant public domain codes for all three types of analysis; (2) evaluate the codes for the accuracy of results and computational efficiency; (3) develop aero-thermal and thermal structural mapping algorithms; (4) integrate all the codes into one single code; (5) write a graphical user interface to improve the user friendliness of the code; (6) conduct test studies for rocket based combined-cycle engine inlet; and finally (7) fabricate a demonstration inlet model using textile preform composites. Tasks one, two and six are being pursued. Selected and evaluated NPARC for flow field analysis, CSTEM for in-depth thermal analysis of inlets and nozzles and FRAC3D for stress analysis. These codes have been independently verified for accuracy and performance. In addition, graphical user interface based on micromechanics analysis for laminated as well as textile composites was developed. Demonstration of this code will be made at the conference. A rocket based combined cycle engine was selected for test studies. Flow field analysis of various inlet geometries were studied. Integration of codes is being continued. The codes developed are being applied to a candidate example of trailblazer engine proposed for space transportation. A successful development of the code will provide a simpler, faster and user-friendly tool for conducting design studies of aircraft and spacecraft engines, applicable in high speed civil transport and space missions.

Shivakumar, Kuwigai↗
NASA NTRS1999

Applications of Parallel Processing in Configuration Analyses

The paper presents the recent progress made towards developing an efficient and user-friendly parallel environment for routine analysis of large CFD problems. The coarse-grain parallel version of the CFL3D Euler/Navier-Stokes analysis code, CFL3Dhp, has been ported onto most available parallel platforms. The CFL3Dhp solution accuracy on these parallel platforms has been verified with the CFL3D sequential analyses. User-friendly pre- and post-processing tools that enable a seamless transfer from sequential to parallel processing have been written. Static load balancing tool for CFL3Dhp analysis has also been implemented for achieving good parallel efficiency. For large problems, load balancing efficiency as high as 95% can be achieved even when large number of processors are used. Linear scalability of the CFL3Dhp code with increasing number of processors has also been shown using a large installed transonic nozzle boattail analysis. To highlight the fast turn-around time of parallel processing, the TCA full configuration in sideslip Navier-Stokes drag polar at supersonic cruise has been obtained in a day. CFL3Dhp is currently being used as a production analysis tool.

Sundaram, Ppchuraman↗
DOE OSTI2024

3D radiated power analysis of JET SPI discharges using the Emis3D forward modeling tool

Abstract Precise values for radiated energy in tokamak disruption experiments are needed to validate disruption mitigation techniques for burning plasma tokamaks like ITER and SPARC. Control room analysis of radiated power ( P rad ) on JET assumes axisymmetry, since fitting 3D radiation structures with limited bolometry coverage is an under-determined problem. In mitigated disruptions, radiation is toroidally asymmetric and 3D, due to fast-growing 3D MHD modes and localized impurity sources. To address this problem, Emis3D adopts a physics motivated forward modeling (‘guess and check’) approach, comparing experimental bolometry data to synthetic data from user-defined radiation structures. Synthetic structures are observed with the Cherab modeling framework and a best fit chosen using a reduced χ 2 statistic. 2D tomographic inversion models are tested, as well as helical flux tubes and 3D MHD simulated structures from JOREK. Two nominally identical pure neon shattered pellet injection (SPI) mitigated discharges in JET are analyzed. 2D tomographic inversions with added toroidal freedom are the best fits in the thermal quench (TQ) and current quench (CQ). In the pre-TQ, 2D reconstructions are statistically the best fits, but are likely over-optimized and do not capture the 3D radiation structure seen in fast camera images. The next-best pre-TQ fits are helical structures that extend towards the high-field side, consistent with an impurity flow under the magnetic nozzle effect also observed in JOREK simulations. Whole-disruption radiated fractions of 0.98 + 0.03 / − 0.29 and 1.01 + 0.02 / − 0.17 are found, suggesting that the stored energy may have been fully mitigated by each SPI, although mitigation efficiencies well below ITER and SPARC requirements for high energy pulses are still within the large uncertainties. Emis3D is also used to validate JOREK SPI simulations, and confirms improvements in matching experiment from changes to impurity modeling. Time-dependent toroidal peaking factors are calculated and discussed.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1990

Solid-liquid staged combustion space boosters

NASA has begun to evaluate solid-liquid hybrid propulsion for launch vehicle booster. A three-phase program was outlined to identify, acquire, and demonstrate technology needed to approximate solid and liquid propulsion state of the art. Aerojet has completed a Phase 1 study and recommends a solid-liquid staged combustion concept in which turbopump fed LO2 is burned with fuel-rich solid propellant effluent in aft-mounted thrust chambers.These reasonably sized thrust chambers are LO2 regeneratively cooled, supplemented with fuel-rich barrier cooling. Turbopumps are driven by the resulting GO2 coolant in an expander-bleed-burnoff cycle. Turbine exhaust pressurizes the LO2 tankage directly, and the excess is bled into supersonic nozzle splitlines, where it combusts with the fuel rich boundary layer. Thrust vector control is enhanced by supersonic nozzle movement on flexseal mounts. Every hybrid solid-liquid concept examined improves booster energy management and launch propellant safety compared to current solid boosters. Solid-liquid staged combustion improves hybrid performance by improving both combustion efficiency and combustion stability, especially important for large boosters. These improvements result from careful fluid management and use of smaller combustors. The study shows NASA safety, reliability, cost, and performance criteria are best met with this concept, wherein simple hardware relies on several separate emerging technologies, all of which have been demonstrated successfully.

Culver, D. W.↗
NASA NTRS1993

A Large Hemi-Anechoic Enclosure for Community-Compatible Aeroacoustic Testing of Aircraft Propulsion Systems

A large hemi-anechoic (absorptive walls and acoustically hard floor) noise control enclosure has been erected around a complex of test stands at the NASA Lewis Research Center in Cleveland, Ohio. This new state-of-the-art Aeroacoustic Propulsion Laboratory (APL) provides an all-weather, semisecure test environment while limiting noise to acceptable levels in surrounding residential neighborhoods. The 39.6 m (130 ft) diameter geodesic dome structure houses the new Nozzle Aeroacoustic Test Rig (NATR), an ejector-powered M = 0.3 free jet facility for acoustic testing of supersonic aircraft exhaust nozzles and turbomachinery. A multi-axis, force-measuring Powered Lift Facility (PLF) stand for testing of Short Takeoff Vertical Landing (STOVL) vehicles is also located within the dome. The design of the Aeroacoustic Propulsion Laboratory efficiently accomodates the research functions of two separate test rigs, one of which (NATR) requires a specialized environment for taking acoustic measurements. Absorptive fiberglass wedge treatment on the interior surface of the dome provides a hemi-anechoic interior environment for obtaining the accurate acoustic measurements required to meet research program goals. The APL is the first known geodesic dome structure to incorporate transmission-loss properties as well as interior absorption into a free-standing, community-compatible, hemi-anechoic test facility.

Cooper, Beth A.↗
NASA NTRS1989

A numerical study of chemically reacting flow in nozzles

The space station uses small rocket motors, called thrusters, for orientation control. Because of the lack of viable design tools for small rockets, the initial thruster design was basically a very small version of a large rocket motor. Thrust measurements of the initial design were lower than predicted. To improve predictions it was decided to develop a verison of the RPLUS2D reacting flow code for thruster calculations. RPLUS2D employs an implicit finite volume, lower-upper symmetric successive overrelaxation (LU-SSOR) scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a coupled and very efficient manner. The combustion processes are modeled by a 9-species, 18 step finite-rate chemistry model, and the turbulence is simulated by a Baldwin-Lomax algebraic model. The code is extended to handle multiple subsonic inlet conditions where the total mass flow is governed by conditions calculated at the thruster-throat. Results are shown for a thruster design where the overall mixture ratio is hydrogen rich. A calculation of a large area ratio divergent nozzle is also presented.

Vanoverbeke, Thomas J.↗
NASA NTRS1989

A numerical study of chemically reacting flow in nozzles

The Space Station uses small rocket motors, called thrusters, for orientation control. Because of the lack of viable design tools for small rockets, the initial thruster design was basically a very small version of a large rocket motor. Thrust measurements of the initial design were lower than predicted. To improve predictions it was decided to develop a version of the RPLUS2D reacting flow code for thruster calculations. RPLUS2D employs an implicit finite volume, lower-upper symmetric successive overrelaxation (LU-SSOR) scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a coupled and very efficient manner. The combustion processes are modeled by a 9-species, 18 step finite-rate chemistry model, and the turbulence is simulated by a Baldwin-Lomax algebraic model. The code is extended to handle multiple subsonic inlet conditions where the total mass flow is governed by conditions calculated at the thruster-throat. Results are shown for a thruster design where the overall mixture ratio is hydrogen rich. A calculation of a large area ratio divergent nozzle is also presented.

Vanoverbeke, Thomas J.↗
DOE OSTI2022

Bulk Spray and Individual Plume Characterization of LPG and Iso-Octane Sprays at Engine-Like Conditions

This study presents experimental and numerical examination of directly injected (DI) propane and iso-octane, surrogates for liquified petroleum gas (LPG) and gasoline, respectively, at various engine like conditions with the overall objective to establish the baseline with regards to fuel delivery required for future high efficiency DI-LPG fueled heavy-duty engines. Sprays for both iso-octane and propane were characterized and the results from the optical diagnostic techniques including high-speed Schlieren and planar Mie scattering imaging were applied to differentiate the liquid-phase regions and the bulk spray phenomenon from single plume behaviors. The experimental results, coupled with high-fidelity internal nozzle-flow simulations were then used to define best practices in CFD Lagrangian spray models. Optical imaging revealed that unlike iso-octane, propane’s spray propagation was fed by its flash boiling, spray collapse, and high degree of vaporization, resulting in a direct proportionality of propane’s penetration length to temperature. These unique features of propane and its variation from iso-octane’s spray pattern, contributed to its classification as an unconventional spray. Appropriate corrections to the injection and breakup models were developed to reproduce the under-expanded jet dynamics and to mimic the flash boiling-driven spray development observed with propane sprays. The simulation results were found to be sensitive to cone and inclusion angles of the blob injector. The current work represents a first assessment of the capability of the commonly available models for engine-spray simulations and highlights the fact that, despite the reasonable agreement obtained in the fuel vapor morphology, the representation of the liquid phase lacks accuracy and requires further model development.

02 PETROLEUM↗
NASA NTRS1993

Advanced Multi-Phase Flow CFD Model Development for Solid Rocket Motor Flowfield Analysis

It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better prediction of the ASRM performance. The multi-phase flow analysis using the FDNS code in the present research can be used as a design tool for solid rocket motor applications.

Liaw, Paul↗
NASA NTRS1931

The Design and Development of an Automatic Injection Valve with an Annular Orifice of Varying Area

The purpose of this investigation was to provide an automatic injection valve of simple construction which would produce a finely atomized oil spray of broad cone angle and would fulfill the requirements of fuel injection in aircraft oil engines. The injection valve designed has only six parts - i. e., two concentric nozzle tubes flared at one end, two body parts, and two nuts. Analysis and engine tests indicate that the fuel spray from this type of injection valve has characteristics which reduce the time lag of autoignition and promote efficient combustion in high-speed oil engines.

Joachim, William F↗