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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Review of the PDWA Concept for Combustion Enhancement in a Supersonic Air-Breathing Combustor Environment

This paper reviews the design of the Pulsed Detonation Wave Augmentor (PDWA) concept and the preliminary computational fluid dynamics studies that supported it. The PDWA relies on the rapid generation of detonation waves in a small tube, which are then injected into the supersonic stream of the main combustor. The blast waves thus generated are used to stimulate the mixing and combustion inside the main combustor. The mixing enhancement relies on various forms of the baroclinic interaction, where misaligned pressure and density gradients combine to produce vortical flow. By using unsteady shock waves, the concept also uses the Richtmyer-Meshkov effect to further increase the rate of mixing. By carefully designing the respective configurations of the combustor and the detonation tubes, one can also increase the penetration of the fuel into the supersonic air stream. The unsteady shocks produce lower stagnation pressure losses than steady shocks. Combustion enhancement can also be obtained through the transient shock-heating of the fuel-air interface, and the lowering of the ignition delay in these regions. The numerical simulations identify these processes, and show which configurations give the best results. Engineering considerations are also presented, and discuss the feasibility of the concept. Of primary importance are the enhancements in performance, the design simplicity, the minimization of the power, cost, and weight, and the methods to achieve very rapid cycling.

Canbier, Jean-Luc↗
NASA NTRS1991

Scramjet mixing establishment times for a pulse facility

A numerical simulation of the temporally developing flow through a generic scramjet combustor duct is presented for stagnation conditions typical of flight at Mach 13 as produced by a shock tunnel pulse facility. The particular focus is to examine the start up transients and to determine the time required for certain flow parameters to become established. The calculations were made with a Navier-Stokes solver SPARK with temporally relaxing inflow conditions derived from operation of the T4 shock tunnel at the University of Queensland in Australia. Calculations at nominal steady inflow conditions were made for comparison. The generic combustor geometry includes the injection of hydrogen fuel from the base of a centrally located strut. In both cases, the flow was assumed laminar and fuel combustion was not included. The establishment process is presented for viscous parameters in the boundary layer and for parameters related to the fuel mixing.

Rogers, R. Clayton↗
NASA NTRS1991

Liquid rocket combustion instability analysis by CFD methods

Combustion instability in liquid rocket engines is simulated computationally by using a simple two-parameter model for the combustion response function. The objectives of the study are to assess the capabilities of CFD algorithms for instability studies and to investigate the response to parametric effects such as bombs and distributed combustion. Results indicate that numerical solutions of high accuracy can be obtained if a sufficient number of grid points are used per wavelength of the disturbance. The short-term response to bombs or pulses triggers a large number of modes in the combustor whose faithful resolution requires highly dense grids, although there is evidence that correct long-term solutions can be obtained even if all the short-term frequencies are not resolved. Long-term responses to pulses are shown to decay to the most unstable mode in small amplitude cases, and to exhibit limit cycles in large amplitude cases. Comparison of distributed with concentrated heat release indicates the former is more stable for given values of the combustion response parameters, and that the distributed heat release gives rise to higher frequency disturbances. Wave steepening is observed in the solutions, but its effect is less pronounced in multidimensional waves than in one-dimensional waves.

Grenda, J. M.↗
NASA NTRS1992

Operating characteristics at Mach 4 of an inlet having forward-swept, sidewall-compression surfaces

A three-dimensional, forward-swept sidewall-compression inlet has been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to examine parameters that affect inlet starting, operability, and performance. The inlet was designed to simultaneously provide good starting and mass capture characteristics through the combination of forward-swept sidewalls and the aft placement of the downstream cowl. Parametrics examined included top surface compression, geometric contraction ratio, sidewall compression angle, boundary-layer thickness on the vehicle undersurface, and cowl position. To simulate combustion effects on inlet operation, the inlet was mechanically back pressured to determine maximum achievable combustor-to-inlet pressure ratio prior to inlet unstart. Inlet operability and performance are shown in terms of inlet starting (pulsed and self-starting) and unstarting characteristics, inlet static pressure distributions, combustor back pressure limits, mass capture, and flow uniformity.

Hudgens, Julie A.↗
NASA NTRS1997

Combustion Noise at Elevated Pressures in a Liquid-Fueled Premixed Combustor

Noise generated in gas turbine combustors can exist in several forms-broadband noise, sharp resonant peaks, and regular or intermittent nonlinear pulsing. In the present study, dynamic pressure measurements were made in several JP-5-fueled combustor configurations, at various mean pressures and temperatures. The fluctuating pressure was measured at mean pressures from 6 to 14 atm and inlet temperatures from 550 K to 850 K. The goal of the present work was to study the effect of changes in mean flow conditions on combustor noise: both broadband noise and sharp tones were considered. In general, the shape of the broadband noise spectrum was consistent from one configuration to another. The shape of the spectrum was influenced by the acoustic filtering of the combustion zone. This filtering ensured the basic consistency of the spectra. In general, the trends in broadband noise observed at low mean pressures were also seen at high mean pressures; that is, the total sound level decreased with both increasing equivalence ratio and increasing inlet temperature. The combustor configurations without a central pilot experienced higher broadband noise levels and were more susceptible to narrow peak resonances than configurations with a central pilot. The sharp peaks were more sensitive to the mean flow than was the broadband noise, and the effects were not always the same. In some situations, increasing the equivalence ratio made the sharp peaks grow, while at other conditions, increasing the equivalence ratio made the sharp peaks shrink. Thus, it was difficult to predict when resonances would occur; however, they were reproducible. Acoustic coupling between the upstream and downstream regions of the combustor may play a role in the sharp-peaked oscillations. Noise was also observed near lean blow out. As with other types of noise, lean blow out noise was affected by the combustion chamber acoustics, which apparently maintains the fluctuations at a uniform frequency. However, the actual conditions when this type of noise was experienced appeared to simply follow the lean blow out limit as it varied with mean temperature and pressure.

Darling, Douglas↗
NASA NTRSDate not supplied

Isolator Shock Train Response to Dual-Mode Scramjet Throttling-Like Forcing

Typical dual-mode scramjet designs employ an isolator component situated between the inlet and combustor. The isolator fluidically buffers the inlet from the combustor via the formation of a shock train, an interconnected series of shock-boundary layer interactions. The isolator is required during ramjet-engine operation and at the lower edges of the scramjet operational envelope. In an operational engine, transient behavior imposes fluidic dynamics due to attitude adjustments, combustor throttling, aberrant resonances within the combustor, etc. To accommodate these transients, the shock train adjusts on both temporal and spatial scales. Sufficiently large downstream pressure increases in the combustor play a critical role during unstart, whereby strong compression waves propagate upstream in the isolator to ultimately force the shock train out the inlet, if it cannot stabilize to the pressure increase. This critically reduces the engine mass flow rate, potentially causing a catastrophic reduction in thrust and loss of the vehicle. A nearly instantaneous unstart can, in principle, be predicted by a ratio of the normal shock pressure jump in an inviscid isolator. However, lower-magnitude perturbations, below this bound, also induce unstarts. Understanding the transient behavior inherent in unsteady isolator operation is critical to reducing safety margin in isolator length and refining engine control algorithms to improve engine performance. The Isolator Dynamics Research Lab (IDRL) is a cold-flow, direct-connect isolator facility ideal for assessing the fundamental fluid dynamics of shock trains. The internal flow of a scramjet isolator is modeled using a 2D, symmetric Mach 2.5 nozzle followed by a 695 mm long, 25.4x50.8 mm rectangular isolator test section. A blockage cone on a high-speed ball screw actuator downstream of the isolator is pulsed to simulate the throttling behavior in an operational combustor. Wall flush high-speed static pressure transducers detect the streamwise shock train within the isolator. Time delays between the onset of the throttling and the response motion of the shock train are analyzed to assess the controlling scales and parameters. From this analysis, the response time of the isolator shock train system is found to depend primarily on the isolator back pressure time-rate-of-change but quickly saturates to a maximum propagation speed along the isolator. These results inform dual-mode scramjet engine designers of the critical parameters and limits of an unstart resistant engine.

scramjet↗
NASA NTRS1988

Concentration, temperature, and density in a hydrogen-air flame by excimer-induced Raman scattering

Single-pulse, vibrational Raman scattering (VRS) is an attractive laser diagnostic for the study of supersonic hydrogen-air combustion. The VRS technique gives a complete thermodynamic description of the gas mixture at a point in the reacting flow. Single-pulse, vibrational Raman scattering can simultaneously provide independent measurements of density, temperature, and concentration of each major species (H2, H2O, O2 and N2) in a hydrogen/air turbulent combustor. Also the pressure can be calculated using the ideal gas law. However, single-pulse VRS systems in current use for measurement of turbulent combustion have a number of shortcomings when applied to supersonic flows: (1) slow repetition rate (1 to 5 Hz), (2) poor spatial resolution (0.5x0.3x0.3 cu mm), and (3) marginal time resolution. Most of these shortcomings are due to the use of visible wavelength flash-lamp pumped dye lasers. The advent of UV excimer laser allows the possibility of dramatic improvements in the single-pulse, vibrational Raman scattering. The excimer based VRS probe will greatly improve repetition rate (100 to 500 Hz), spatial resolution (0.1x0.1x0.1 cu mm) and time resolution (30ns). These improvements result from the lower divergence of the UV excimer, higher repetition rate, and the increased Raman cross-sections (15 to 20 times higher) at ultra-violet (UV) wavelengths. With this increased capability, single-pulse vibrational Raman scattering promises to be an ideal non-intrusive probe for the study of hypersonic propulsion flows.

Wehrmeyer, Joseph A.↗
NASA NTRS1999

An Experiment Investigation of Fully-Modulated, Turbulent Diffusion Flames in Reduced Gravity

Pulsed combustion appears to have the potential to provide for rapid fuel/air mixing, compact and economical combustors, and reduced exhaust emissions. The ultimate objective of this program is to increase the fundamental understanding of the fuel/air mixing and combustion behavior of pulsed, turbulent diffusion flames by conducting experiments in microgravity. In this research the fuel jet is fully-modulated (i.e., completely shut off between pulses) by an externally controlled valve system. This can give rise to drastic modification of the combustion and flow characteristics of flames, leading to enhanced fuel/air mixing mechanisms not operative for the case of acoustically excited or partially-modulated jets. In addition, the fully-modulated injection approach avoids the strong acoustic forcing present in pulsed combustion devices, significantly simplifying the mixing and combustion processes. Relatively little is known of the behavior of turbulent flames in reduced-gravity conditions, even in the absence of pulsing. The goal of this Flight-Definition experiment (PUFF, for PUlsed-Fully Flames) is to establish the behavior of fully-modulated, turbulent diffusion flames under microgravity conditions. Fundamental issues to be addressed in this experiment include the mechanisms responsible for the flame length decrease for fully-modulated, turbulent diffusion flames compared with steady flames, the impact of buoyancy on the mixing and combustion characteristics of these flames, and the characteristics of turbulent flame puffs under fully momentum-dominated conditions.

Hermanson, J. C.↗
NASA NTRS1986

Detonation wave compression in gas turbines

A study was made of the concept of augmenting the performance of low pressure ratio gas turbines by detonation wave compression of part of the flow. The concept exploits the constant volume heat release of detonation waves to increase the efficiency of the Brayton cycle. In the models studied, a fraction of the compressor output was channeled into detonation ducts where it was processed by transient transverse detonation waves. Gas dynamic studies determined the maximum cycling frequency of detonation ducts, proved that upstream propagation of pressure pulses represented no problems and determined the variations of detonation duct output with time. Mixing and wave compression were used to recombine the combustor and detonation duct flows and a concept for a spiral collector to further smooth the pressure and temperature pulses was presented as an optional component. The best performance was obtained with a single firing of the ducts so that the flow could be re-established before the next detonation was initiated. At the optimum conditions of maximum frequency of the detonation ducts, the gas turbine efficiency was found to be 45 percent while that of a corresponding pressure ratio 5 conventional gas turbine was only 26%. Comparable improvements in specific fuel consumption data were found for gas turbines operating as jet engines, turbofans, and shaft output machines. Direct use of the detonation duct output for jet propulsion proved unsatisfactory. Careful analysis of the models of the fluid flow phenomena led to the conclusion that even more elaborate calculations would not diminish the uncertainties in the analysis of the system. Feasibility of the concept to work as an engine now requires validation in an engineering laboratory experiment.

Wortman, A.↗
NASA NTRS1997

Fuel Line Based Acoustic Flame-Out Detection System

An acoustic flame-out detection system that renders a large high pressure combustor safe in the event of a flame-out and possible explosive reignition. A dynamic pressure transducer is placed in the fuel and detects the stabilizing fuel pressure oscillations, caused by the combustion process. An electric circuit converts the signal from the combustion vortices, and transmitted to the fuel flow to a series of pulses. A missing pulse detector counts the pulses and continuously resets itself. If three consecutive pulses are missing, the circuit closes the fuel valve. With fuel denied the combustor is shut down or restarted under controlled conditions.

Puster, Richard L.↗
NASA NTRS2002

Active Control of Combustor Instability Shown to Help Lower Emissions

In a quest to reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities, or high-pressure oscillations much like sound waves, that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Aerospace Propulsion and Power Base Research and Technology Program, the NASA Glenn Research Center, in partnership with Pratt & Whitney and United Technologies Research Center, is developing technologies for the active control of combustion instabilities. With active combustion control, the fuel is pulsed to put pressure oscillations into the system. This cancels out the pressure oscillations being produced by the instabilities. Thus, the engine can have lower pollutant emissions and long life.The use of active combustion instability control to reduce thermo-acoustic-driven combustor pressure oscillations was demonstrated on a single-nozzle combustor rig at United Technologies. This rig has many of the complexities of a real engine combustor (i.e., an actual fuel nozzle and swirler, dilution cooling, etc.). Control was demonstrated through modeling, developing, and testing a fuel-delivery system able to the 280-Hz instability frequency. The preceding figure shows the capability of this system to provide high-frequency fuel modulations. Because of the high-shear contrarotating airflow in the fuel injector, there was some concern that the fuel pulses would be attenuated to the point where they would not be effective for control. Testing in the combustor rig showed that open-loop pulsing of the fuel was, in fact, able to effectively modulate the combustor pressure. To suppress the combustor pressure oscillations due to thermoacoustic instabilities, it is desirable to time the injection of the fuel so that it interferes with the instability. A closed-loop control scheme was developed that uses combustion pressure feedback and a phase-shifting controller to time the fuel-injection pulses. Some suppression of the pressure oscillations at the 280-Hz instability frequency was demonstrated (see the next figure). However, the overall peak-to- peak pressure oscillations in the combustor were only mildly reduced. Improvements to control hardware and control methods are being continued to gain improved closed-loop reduction of the pressure oscillations.pulse the fuel at

DeLaat, John C.↗
NASA NTRS1992

Experimental Investigation of Nozzle/Plume Aerodynamics at Hypersonic Speeds

The work performed by D. W. Bogdanoff and J.-L. Cambier during the period of 1 Feb. - 31 Oct. 1992 is presented. The following topics are discussed: (1) improvement in the operation of the facility; (2) the wedge model; (3) calibration of the new test section; (4) combustor model; (5) hydrogen fuel system for combustor model; (6) three inch calibration/development tunnel; (7) shock tunnel unsteady flow; (8) pulse detonation wave engine; (9) DCAF flow simulation; (10) high temperature shock layer simulation; and (11) the one dimensional Godunov CFD code.

Heinemann, K.↗
NASA NTRS2002

Oxidation- and Creep-Enhanced Fatigue of Haynes 188 Alloy-Oxide Scale System Under Simulated Pulse Detonation Engine Conditions

The development of the pulse detonation engine (PDE) requires robust design of the engine components that are capable of enduring harsh detonation environments. In this study, a high cycle thermal fatigue test rig was developed for evaluating candidate PDE combustor materials using a CO2 laser. The high cycle thermal fatigue behavior of Haynes 188 alloy was investigated under an enhanced pulsed laser test condition of 30 Hz cycle frequency (33 ms pulse period, and 10 ms pulse width including 0.2 ms pulse spike). The temperature swings generated by the laser pulses near the specimen surface were characterized by using one-dimensional finite difference modeling combined with experimental measurements. The temperature swings resulted in significant thermal cyclic stresses in the oxide scale/alloy system, and induced extensive surface cracking. Striations of various sizes were observed at the cracked surfaces and oxide/alloy interfaces under the cyclic stresses. The test results indicated that oxidation and creep-enhanced fatigue at the oxide scale/alloy interface was an important mechanism for the surface crack initiation and propagation under the simulated PDE condition.

Zhu, Dongming↗
NASA NTRS1983

Atomization and combustion performance of antimisting kerosene and jet fuel

Combustion performance of antimisting kerosene (AMK) containing FM-9 polymer was investigated at various levels of degradation (restoration of AMK for normal use in a gas turbine engine). To establish the relationship of degradation and atomization to performance in an aircraft gas turbine combustor, sprays formed by the nozzle of a JT8-D combustor with Jet A and AMK at 1 atmosphere (atm) (14.1 lb/square in absolute) pressure and 22 C at several degradation levels were analyzed. A new spray characterization technique based on digital image analysis of high resolution, wide field spray images formed under pulsed ruby laser sheet illumination was developed. Combustion tests were performed for these fuels in a JT8-D single can combustor facility to measure combustion efficiency and the lean extinction limit. Correlation of combustion performance under simulated engine operating conditions with nozzle spray Sauter mean diameter (SMD) measured at 1 atm and 22 C were observed. Fuel spray SMD and hence the combustion efficiency are strongly influenced by fuel degradation level. Use of even the most highly degraded AMK tested (filter ratio = 1.2) resulted in an increase in fuel consumption of 0.08% to 0.20% at engine cruise conditions.

Fleeter, R.↗
NASA NTRS1993

Axisymmetric single shear element combustion instability experiment

The combustion stability characteristics of a combustor consisting of a single shear element and a cylindrical chamber utilizing LOX and gaseous hydrogen as propellants are presented. The combustor geometry and the resulting longitudinal mode instability are axisymmetric. Hydrogen injection temperature and pyrotechnic pulsing were used to determine stability boundaries. Mixture ratio, fuel annulus gap, and LOX post configuration were varied. Performance and stability data are presented for chamber pressures of 300 and 1000 psia.

Breisacher, Kevin J.↗
NASA NTRS1993

Axisymmetric single shear element combustion instability experiment

The combustion stability characteristics of a combustor consisting of a single shear element and a cylindrical chamber utilizing LOX and gaseous hydrogen as propellants are presented. The combustor geometry and the resulting longitudinal mode instability are axisymmetric. Hydrogen injection temperature and pyrotechnic pulsing were used to determine stability boundaries. Mixture ratio, fuel annulus gap, and LOX post configuration were varied. Performance and stability data were obtained for chamber pressures of 300 and 1000 psia.

Breisacher, Kevin J.↗
NASA NTRS1994

Experimental investigation of nozzle/plume aerodynamics at hypersonic speeds

Much of the work involved the Ames 16-Inch Shock Tunnel facility. The facility was reactivated and upgraded, a data acquisition system was configured and upgraded several times, several facility calibrations were performed and test entries with a wedge model with hydrogen injection and a full scramjet combustor model, with hydrogen injection, were performed. Extensive CFD modeling of the flow in the facility was done. This includes modeling of the unsteady flow in the driver and driven tubes and steady flow modeling of the nozzle flow. Other modeling efforts include simulations of non-equilibrium flows and turbulence, plasmas, light gas guns and the use of non-ideal gas equations of state. New experimental techniques to improve the performance of gas guns, shock tubes and tunnels and scramjet combustors were conceived and studied computationally. Ways to improve scramjet engine performance using steady and pulsed detonation waves were also studied computationally. A number of studies were performed on the operation of the ram accelerator, including investigations of in-tube gasdynamic heating and the use of high explosives to raise the velocity capability of the device.

Bogdanoff, David W.↗
NASA NTRS2002

Pulse Detonation Engine Test Bed Developed

A detonation is a supersonic combustion wave. A Pulse Detonation Engine (PDE) repetitively creates a series of detonation waves to take advantage of rapid burning and high peak pressures to efficiently produce thrust. NASA Glenn Research Center's Combustion Branch has developed a PDE test bed that can reproduce the operating conditions that might be encountered in an actual engine. It allows the rapid and cost-efficient evaluation of the technical issues and technologies associated with these engines. The test bed is modular in design. It consists of various length sections of both 2- and 2.6- in. internal-diameter combustor tubes. These tubes can be bolted together to create a variety of combustor configurations. A series of bosses allow instrumentation to be inserted on the tubes. Dynamic pressure sensors and heat flux gauges have been used to characterize the performance of the test bed. The PDE test bed is designed to utilize an existing calorimeter (for heat load measurement) and windowed (for optical access) combustor sections. It uses hydrogen as the fuel, and oxygen and nitrogen are mixed to simulate air. An electronic controller is used to open the hydrogen and air valves (or a continuous flow of air is used) and to fire the spark at the appropriate times. Scheduled tests on the test bed include an evaluation of the pumping ability of the train of detonation waves for use in an ejector and an evaluation of the pollutants formed in a PDE combustor. Glenn's Combustion Branch uses the National Combustor Code (NCC) to perform numerical analyses of PDE's as well as to evaluate alternative detonative combustion devices. Pulse Detonation Engine testbed.

Breisacher, Kevin J.↗