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NASA NTRS2002

Dual-Mode Combustion

The Department of Mechanical and Aerospace Engineering at the University of Virginia has conducted an investigation of the mixing and combustion processes in a hydrogen fueled dual-mode scramjet combustor. The experiment essentially consisted of the "direct connect" continuous operation of a Mach 2 rectangular combustor with a single unswept ramp fuel injector. The stagnation enthalpy of the test flow simulated a flight Mach number of 5. Measurements were obtained using conventional wall instrumentation and laser based diagnostics. These diagnostics included, pressure and wall temperature measurements, Fuel Plume Imaging (FPI) and Particle Image Velocimetry (PIV). A schematic of the combustor configuration and a summary of the measurements obtained are presented. The experimental work at UVa was parallel by Computational Fluid Dynamics (CFD) work at NASA Langley. The numerical and experiment results are compared in this document.

Goyne, Christopher P.↗

NASA NTRS1990

Approach to SSME health monitoring. II - Exhaust plume emission spectroscopy at the DTF

The Diagnostics Testbed Facility (DTF) located at the Stennis Space Center (SSC) is used for obtaining extensive sets of H2O2 exhaust plume emission spectral data for the SSME critical components related elements and materials. The SSME related elements and materials are simulated by mixing appropriate amounts of compounds of their respective constituent elements in an aqueous solution which is injected into the combustion chamber of the DTFT. Five of the most critical components of the SSME which have experienced very severe wear and tear problems in the past are analyzed. These are high pressure turbopump (HPTP) turbine blades, HPTP turbine disks, HPTP bearing, main injector LOX posts, and the main combustion chamber structural shell. The alloys used in the manufacturing of these components are MAR-M 246 + Hf, Waspaloy X, AISI 440C, Haynes 188, and Inconel 718, respectively. The experimental setup and procedures at the DTF are described; stratospheric data for the five alloys are presented; and strategies for the material identification in the SSME exhaust plume are discussed.

Tejwani, Gopal D.↗

NASA NTRS2001

Microgravity Combustion Science and Fluid Physics Experiments and Facilities for the ISS

At the NASA Glenn Research Center, the Microgravity Science Program supports both ground-based and flight experiment research in the disciplines of Combustion Science and Fluid Physics. Combustion Science research includes the areas of gas jet diffusion flames, laminar flames, burning of droplets and misting fuels, solids and materials flammability, fire and fire suppressants, turbulent combustion, reaction kinetics, materials synthesis, and other combustion systems. The Fluid Physics discipline includes the areas of complex fluids (colloids, gels, foams, magneto-rheological fluids, non-Newtonian fluids, suspensions, granular materials), dynamics and instabilities (bubble and drop dynamics, magneto/electrohydrodynamics, electrochemical transport, geophysical flows), interfacial phenomena (wetting, capillarity, contact line hydrodynamics), and multiphase flows and phase changes (boiling and condensation, heat transfer, flow instabilities). A specialized International Space Station (ISS) facility that provides sophisticated research capabilities for these disciplines is the Fluids and Combustion Facility (FCF). The FCF consists of the Combustion Integrated Rack (CIR), the Fluids Integrated Rack (FIR) and the Shared Accommodations Rack and is designed to accomplish a large number of science investigations over the life of the ISS. The modular, multiuser facility is designed to optimize the science return within the available resources of on-orbit power, uplink/downlink capacity, crew time, upmass/downmass, volume, etc. A suite of diagnostics capabilities, with emphasis on optical techniques, will be provided to complement the capabilities of the subsystem multiuser or principal investigator-specific experiment modules. The paper will discuss the systems concept, technical capabilities, functionality, and the initial science investigations in each discipline.

Lauver, Richard W.↗

NASA NTRS1990

Progress in hypersonic combustion technology with computation and experiment

Design of successful airbreathing engines for operation at near-orbital speeds presents significant challenges in all the disciplines involved, including propulsion. This paper presents a discussion of the important physics of hypersonic combustion and an assessment of the state of the art of ground simulations with pulse facilities and with computational techniques. Recent examples of experimental and computational simulations are presented and discussed. The need for continued application of these tools to establish the credibility and fidelity of engineering design methods for practical hypersonic combustors is emphasized along with the critical need for improved diagnostic methods for hypervelocity reacting flows.

Anderson, Griffin Y.↗

NASA NTRS1990

CARS Temperature Measurements in a Hypersonic Propulsion Test Facility

Nonintrusive diagnostic measurements were performed in the supersonic reacting flow of the Hypersonic Propulsion Test Cell 2 at NASA-Langley. A Coherent Anti-stokes Raman Spectroscopy (CARS) system was assembled specifically for the test cell environment. System design considerations were: (1) test cell noise and vibration; (2) contamination from flow field or atmospheric borne dust; (3) unwanted laser or electrically induced combustion (inside or outside the duct); (4) efficient signal collection; (5) signal splitting to span the wide dynamic range present throughout the flow field; (6) movement of the sampling volume in the flow; and (7) modification of the scramjet model duct to permit optical access to the reacting flow with the CARS system. The flow in the duct was a nominal Mach 2 flow with static pressure near one atmosphere. A single perpendicular injector introduced hydrogen into the flow behind a rearward facing step. CARS data was obtained in three planes downstream of the injection region. At least 20 CARS data points were collected at each of the regularly spaced sampling locations in each data plane. Contour plots of scramjet combustor static temperature in a reacting flow region are presented.

Jarrett, Olin, Jr.↗

NASA NTRS1993

Intelligent neuroprocessors for in-situ launch vehicle propulsion systems health management

Efficacy of existing on-board propulsion systems health management systems (HMS) are severely impacted by computational limitations (e.g., low sampling rates); paradigmatic limitations (e.g., low-fidelity logic/parameter redlining only, false alarms due to noisy/corrupted sensor signatures, preprogrammed diagnostics only); and telemetry bandwidth limitations on space/ground interactions. Ultra-compact/light, adaptive neural networks with massively parallel, asynchronous, fast reconfigurable and fault-tolerant information processing properties have already demonstrated significant potential for inflight diagnostic analyses and resource allocation with reduced ground dependence. In particular, they can automatically exploit correlation effects across multiple sensor streams (plume analyzer, flow meters, vibration detectors, etc.) so as to detect anomaly signatures that cannot be determined from the exploitation of single sensor. Furthermore, neural networks have already demonstrated the potential for impacting real-time fault recovery in vehicle subsystems by adaptively regulating combustion mixture/power subsystems and optimizing resource utilization under degraded conditions. A class of high-performance neuroprocessors, developed at JPL, that have demonstrated potential for next-generation HMS for a family of space transportation vehicles envisioned for the next few decades, including HLLV, NLS, and space shuttle is presented. Of fundamental interest are intelligent neuroprocessors for real-time plume analysis, optimizing combustion mixture-ratio, and feedback to hydraulic, pneumatic control systems. This class includes concurrently asynchronous reprogrammable, nonvolatile, analog neural processors with high speed, high bandwidth electronic/optical I/O interfaced, with special emphasis on NASA's unique requirements in terms of performance, reliability, ultra-high density ultra-compactness, ultra-light weight devices, radiation hardened devices, power stringency, and long life terms.

Gulati, S.↗

NASA NTRS1993

Apparatus for testing high pressure injector elements

An apparatus for testing and evaluating the spray pattern of high pressure fuel injector elements for use in supplying fuel to combustion engines is presented. Prior art fuel injector elements were normally tested by use of low pressure apparatuses which did not provide a purge to prevent mist from obscuring the injector element or to prevent frosting of the view windows; could utilize only one fluid during each test; and had their viewing ports positioned one hundred eighty (180 deg) apart, thus preventing optimum use of laser diagnostics. The high pressure fluid injector test apparatus includes an upper hub, an upper weldment or housing, a first clamp and stud/nut assembly for securing the upper hub to the upper weldment, a standoff assembly within the upper weldment, a pair of window housings having view glasses within the upper weldment, an injector block assembly and purge plate within the upper weldment for holding an injector element to be tested and evaluated, a lower weldment or housing, a second clamp and stud/nut assembly for securing the lower weldment to the upper weldment, a lower hub, a third clamp and stud/nut assembly for securing the lower hub to the lower weldment, mechanisms for introducing fluid under high pressure for testing an injector element, and mechanisms for purging the apparatus to prevent frosting of view glasses within the window housings and to permit unobstructed viewing of the injector element.

Myers, William Neill↗

NASA NTRS1995

Apparatus for testing high pressure injector elements

An apparatus for testing and evaluating the spray pattern of high pressure fuel injector elements for use in supplying fuel to combustion engines is presented. Prior art fuel injector elements were normally tested by use of low pressure apparatuses which did not provide a purge to prevent mist from obscuring the injector element or to prevent frosting of the view windows; could utilize only one fluid during each test; and had their viewing ports positioned one hundred eighty (180 deg) apart, thus preventing optimum use of laser diagnostics. The high pressure fluid injector test apparatus includes an upper hub, an upper weldment or housing, a first clamp and stud/nut assembly for securing the upper hub to the upper weldment, a standoff assembly within the upper weldment, a pair of window housings having view glasses within the upper weldment, an injector block assembly and purge plate within the upper weldment for holding an injector element to be tested and evaluated, a lower weldment or housing, a second clamp and stud/nut assembly for securing the lower weldment to the upper hub, a third clamp and stud/nut assembly for securing the lower hub to the lower weldment, mechanisms for introducing fluid under high pressure for testing an injector element, and mechanisms for purging the apparatus to prevent frosting of view glasses within the window housings and to permit unobstructed viewing of the injector element.

Myers, William Neill↗

NASA NTRS1995

Influence of Geometry and Flow Variations on NO Formation in the Quick Mixer of a Staged Combustor

Staged combustion, such as Rich-Burn/Quick-Mix/Lean-Burn (RQL), is a viable strategy to meet nitric oxide (NO) emission goals for both stationary and propulsion gas turbine engines. A critical element of the design is the quick mixer section where the potential for NO production is high. While numerical calculations of the quick mixer under reacting conditions have been conducted, the hostile environment and lack of appropriate diagnostics have, to date, precluded experimental probing of the reacting case. As an alternative to understanding the effect of geometry and flow variations on the production of NO in the quick mixer, the present paper presents (1) a series of non-reacting parametric studies, and (2) a computational method to extrapolate the results of the non-reacting experiments to reacting conditions. The results show that the rate of NO production is highest in the immediate vicinity of the injection plane. For a given momentum flux ratio between the jets and mainstream, the most effective mixing geometry is that which mixes effectively in both (1) the plane of injection, and (2) the wall regions downstream of the plan of injection. The tailoring of the mixing is key to minimize the NO formed. As a result, the best overall mixer with respect to the minimization of NO production may depend on the system specific characteristics of the particular application.

Hatch, M. S.↗

NASA NTRS1992

Laser absorption flow diagnostics for use in hypersonic ground-based and flight experiments

Consideration is given to laboratory ground-based scale systems for measuring OH and NO and diode laser systems for measuring O2 and H2O. Several diagnostic instruments which are currently being used in a scramjet combustor test rig of the Ames 16-inch shock tunnel are described. Data on mole fraction, temperature, density, pressure, and velocity is obtained using specific data reduction techniques. These measurements provide critical data for evaluation of the propulsion system performance which include mass capture for quantification of fuel equivalence ratio (O2), flow contamination and emission levels (NO), combustion progress (OH), and injector mixing and combustion efficiency (H2O).

Cavolowsky, John A.↗

NASA NTRS2002

Spontaneous Raman Scattering Diagnostics for High-pressure Gaseous Flames

A high-pressure (up to 60 atm) gaseous burner facility with optical access that provides steady, reproducible flames with high precision, and the ability to use multiple fuel/oxidizer combinations has been developed. In addition, a high-performance spontaneous Raman scattering system for use in the above facility has also been developed. Together, the two systems will be used to acquire and establish a comprehensive Raman scattering spectral database for use as a quantitative high-pressure calibration of single-shot Raman scattering measurements in high-pressure combustion systems. Using these facilities, the Raman spectra of H2-Air flames were successfully measured at pressures up to 20 atm. The spectra demonstrated clear rotational and ro-vibrational Raman features of H2, N2, and H2O. theoretical Raman spectra of pure rotational H2, vibrational H2, and vibrational N2 were calculated using a classical harmonic-oscillator model with pressure broadening effects and fitted to the data. At a gas temperature of 1889 K for a phi = 1.34 H2-Air flame, the model and the data showed good agreement, confirming a ro-vibrational equilibrium temperature.

Kojima, Jun↗

NASA NTRS2002

Combustion Sensors: Gas Turbine Applications

This report documents efforts to survey the current research directions in sensor technology for gas turbine systems. The work is driven by the current and future requirements on system performance and optimization. Accurate real time measurements of velocities, pressure, temperatures, and species concentrations will be required for objectives such as combustion instability attenuation, pollutant reduction, engine health management, exhaust profile control via active control, etc. Changing combustor conditions - engine aging, flow path slagging, or rapid maneuvering - will require adaptive responses; the effectiveness of such will be only as good as the dynamic information available for processing. All of these issues point toward the importance of continued sensor development. For adequate control of the combustion process, sensor data must include information about the above mentioned quantities along with equivalence ratios and radical concentrations, and also include both temporal and spatial velocity resolution. Ultimately these devices must transfer from the laboratory to field installations, and thus must become low weight and cost, reliable and maintainable. A primary conclusion from this study is that the optics-based sensor science will be the primary diagnostic in future gas turbine technologies.

Human, Mel↗

NASA NTRS1979

Investigation of the feasibility of CARS measurements in scramjet combustion

Results are presented of analytical and experimental investigations to determine the feasibility of using coherent anti-Stokes Raman spectroscopy (CARS) to measure temperature and species concentrations in supersonic combustion experiments. Experimental CARS spectra of O2 and H2 in a laboratory flat H2 - air diffusion flame are presented. Temperatures deduced from both species generally agree with radiation corrected thermocouple measurements. Oxygen concentrations were determined from the shape of the experimental spectra. Measured H2 CARS signal magnitudes are in good agreement with analytical predictions based on H2 concentration profiles determined from spontaneous Raman scattering. Based on these results, CARS diagnostics appear to be feasible if problems associated with the supersonic flow are not too severe.

Shirley, J. A.↗

NASA NTRS1988

Spacecraft Fire-Safety Experiments for Space Station: Technology Development Mission

Definitions, overall descriptions, and preliminary apparatus and instrumentation requirements are presented for three spacecraft fire-safety experiments appropriate for inclusion in a growth-version of the Space Station. The experiments are as follows: 1. Combustion and Flame Spread of Typical Spacecraft Materials in a Low Velocity Convective Flow in Low Gravity 2. Fire and Fire-Extinguishant Interactions with Various Fire Scenarios in Low Gravity 3. Smoldering and Deep-Seated Combustion in Low Gravity. The experiments were selected from an assessment of recommendations for advanced technologies for prevention, detection, and control of fires in spacecraft. The experiments and their facilities are intended to constitute a portion of a Spacecraft Fire Safety Technology Development Mission (TDM). Basic requirements for the mission study are that the Space Station is essential for the accomplishment of the experimental objectives and that the technology being developed is appropriate for the growth version of the Space Station. The advantages of these three experiments are that, not only do they explore problems relevant to the Space Station operation and utilization, but also they constitute tests adaptable to two multiuse facilities installed in the Space Station laboratory module, the Combustion Tunnel Facility and the Combustion Facility. Three additional spacecraft fire-safety experiments, essential precursors to the TD M for near-term performance, are also described as follows: 1. Near-Term Testing of Materials for Combustion and Flame Spread 2. Near-Term Extinguishment of Fires in a Low-Gravity Test Bed 3. Post-Fire Recovery of Sensitive Electrical and Electronic Equipment. For all the experiments, the study defines accommodations in terms of first-order space, mass, power, consumables, and crew-time estimates. The study also reviews the key problems to be addressed, namely experiment safety hazards, non-intrusive diagnostics development and waste-products disposal. A preliminary task and time schedule for design, development, and experiment fabrication is proposed.

Youngblood, Wallace W.↗

NASA NTRS2007

Radiative Extinction of Gaseous Spherical Diffusion Flames in Microgravity

Radiative extinction of spherical diffusion flames was investigated experimentally and numerically. The experiments involved microgravity spherical diffusion flames burning ethylene and propane at 0.98 bar. Both normal (fuel flowing into oxidizer) and inverse (oxidizer flowing into fuel) flames were studied, with nitrogen supplied to either the fuel or the oxygen. Flame conditions were chosen to ensure that the flames extinguished within the 2.2 s of available test time; thus extinction occurred during unsteady flame conditions. Diagnostics included color video and thin-filament pyrometry. The computations, which simulated flow from a porous sphere into a quiescent environment, included detailed chemistry, transport and radiation, and yielded transient results. Radiative extinction was observed experimentally and simulated numerically. Extinction time, peak temperature, and radiative loss fraction were found to be independent of flow rate except at very low flow rates. Radiative heat loss was dominated by the combustion products downstream of the flame and was found to scale with flame surface area, not volume. For large transient flames the heat release rate also scaled with surface area and thus the radiative loss fraction was largely independent of flow rate. Peak temperatures at extinction onset were about 1100 K, which is significantly lower than for kinetic extinction. One observation of this work is that while radiative heat losses can drive transient extinction, this is not because radiative losses are increasing with time (flame size) but rather because the heat release rate is falling off as the temperature drops.

Santa, K. J.↗

NASA NTRS1987

Automated Reduction of Data from Images and Holograms

Laser techniques are widely used for the diagnostics of aerodynamic flow and particle fields. The storage capability of holograms has made this technique an even more powerful. Over 60 researchers in the field of holography, particle sizing and image processing convened to discuss these topics. The research program of ten government laboratories, several universities, industry and foreign countries were presented. A number of papers on holographic interferometry with applications to fluid mechanics were given. Several papers on combustion and particle sizing, speckle velocimetry and speckle interferometry were given. A session on image processing and automated fringe data reduction techniques and the type of facilities for fringe reduction was held.

Lee, G.↗

NASA NTRS2016

Impact of Flight Enthalpy, Fuel Simulant, and Chemical Reactions on the Mixing Characteristics of Several Injectors at Hypervelocity Flow Conditions

The high total temperatures or total enthalpies required to duplicate the high-speed flight conditions in ground experiments often place stringent requirements on the material selection and cooling needs for the test articles and intrusive flow diagnostic equipment. Furthermore, for internal flows, these conditions often complicate the use of nonintrusive diagnostics that need optical access to the test section and interior portions of the flowpath. Because of the technical challenges and increased costs associated with experimentation at high values of total enthalpy, an attempt is often made to reduce it. This is the case for the Enhanced Injection and Mixing Project (EIMP) currently underway in the Arc-Heated Scramjet Test Facility at the NASA Langley Research Center. The EIMP aims to investigate supersonic combustion ramjet (scramjet) fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships between mixing performance and losses relevant to flight Mach numbers greater than 8. The experiments will consider a "direct-connect" approach and utilize a Mach 6 nozzle to simulate the combustor entrance flow of a scramjet engine. However, while the value of the Mach number is matched to that expected at the combustor entrance in flight, the maximum value of the total enthalpy for these experiments is limited by the thermal-structural limits of the uncooled experimental hardware. Furthermore, the fuel simulant is helium, not hydrogen. The use of "cold" flows and non-reacting mixtures of fuel simulants for mixing experiments is not new and has been extensively utilized as a screening technique for scramjet fuel injectors. In this study, Reynolds-averaged simulations are utilized (RAS) to systematically verify the implicit assumptions used by the EIMP. This is accomplished by first performing RAS of mixing for two injector configurations at planned nominal experimental conditions. The mixing parameters of interest, such as mixing efficiency and total pressure recovery, are then computed and compared to the values obtained from RAS under the true enthalpy conditions and using helium and hydrogen. Finally, the impact of combustion on mixing, often deemed small enough to neglect at hypervelocity conditions, is assessed by comparing the results obtained from the hydrogen-fueled reacting and non-reacting RAS. For reacting flows, in addition to mixing efficiency and total pressure recovery, the combustion efficiency and thrust potential are also considered. In all of the simulations, the incoming air Mach number and the fuel-to-air ratio are the same, while the total pressure, total enthalpy, and the fuel simulant vary depending on the case considered. It is found that under some conditions the "cold" flow experiments are a good approximation of the flight.

Drozda, Tomasz G.↗

NASA NTRS1997

Single-Lever Power Control for General Aviation Aircraft Promises Improved Efficiency and Simplified Pilot Controls

General aviation research is leading to major advances in internal combustion engine control systems for single-engine, single-pilot aircraft. These advances promise to increase engine performance and fuel efficiency while substantially reducing pilot workload and increasing flight safety. One such advance is a single-lever power control (SLPC) system, a welcome departure from older, less user-friendly, multilever engine control systems. The benefits of using single-lever power controls for general aviation aircraft are improved flight safety through advanced engine diagnostics, simplified powerplant operations, increased time between overhauls, and cost-effective technology (extends fuel burn and reduces overhaul costs). The single-lever concept has proven to be so effective in preliminary studies that general aviation manufacturers are making plans to retrofit current aircraft with the technology and are incorporating it in designs for future aircraft.

Musgrave, Jeffrey L.↗