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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
DOE OSTI2022

Optimum Injector Parameters for Thermoacoustic Stability in a Multi-Nozzle Can Combustion System

Abstract High-frequency transverse instabilities are an important concern in can combustor configurations. In these configurations which are typically operated with multiple injectors around a central injector, each injector is subjected to different parts of the acoustic mode shape and thus respond differently for the same instability mode. Recent work by the author has modeled the response of premixed flames to excitation by natural high-frequency transverse modes in a can combustor both in the center and outer nozzles. The stability of these acoustically non-compact flames was assessed using the Rayleigh criterion (Rayleigh Integral denoted as RI) and not the overall unsteady heat release as is the case for compact flames. Several key control parameters were studied, namely — flame angle, swirling strength, nozzle location. For non-axisymmetric modes such as the commonly occurring 1-T mode, both radial and azimuthal offsets of the nozzle location affected stability. The framework was applied to an optimization study to identify the optimal combination of parameters that minimizes RI for the different nozzles in the multi-nozzle system. In this study, a N-around-1 configuration was studied, and the results indicated that the different nozzles needed to be operated at different flame angles and swirl numbers to result in an overall minimum RI. However, the specific response of the different injectors was not considered. The helical mode distribution at each injector varies as we azimuthally go around the combustor’s injector distribution and thus the most amplified mode and the resulting flame response would be different. To minimize RI, it is important to determine the injector configurations that result in a hydrodynamic profile that minimizes the individual RI for each nozzle. The resulting relationship between the injector’s flow and local hydrodynamics can then be used in a hydrodynamics study of an individual injector so that the most optimal injector is chosen depending on its location in the combustor dump plane.

Acharya, Vishal↗
NASA NTRS1977

Update on development of the Space Shuttle main engine /as of 5 May 1977/

The development and testing of the Space Shuttle main engine are described. The preburners, main injector, main combustion chamber, and 35:1 expansion ratio test nozzle were successfully run at full power level (109 percent of rated power level). Integral combustion stability aids damped induced instability oscillations within 6 milliseconds. Three significant turbomachinery problems were identified and solved during the past year. These problems involved high-pressure fuel turbopump subsynchronous whirl, severe overheating of the turbine bearing and components and, in the high-pressure oxidizer turbopump, burning that first occurred in the drain cavity downstream of the primary oxidizer seal.

Johnson, J. R.↗
NASA NTRSDate not supplied

Average Heat Flux Measurements in a Gaseous Detonation-Based Rocket Engine

Rotating detonation rocket engines (RDREs) offer benefits over traditional deflagrative engines, including increased engine performance, compact combustion and negligible detrimental thermoacoustic instabilities due to mode-locking. Realizing the benefits of a detonation-based propulsion system will be accompanied by increased thermal loads due to a combination of heat release at elevated pressure and temperature, as well as compact heat release due to detonation. Managing these heat loads necessitates the development of thermal management strategies. As a first step towards establishing the requirements for detonation-based engine thermal management strategies, this work develops an additively manufactured (AM) water-based, axially resolved calorimeter for integration in an existing 76.8 mm outer diameter (OD) RDRE to measure average chamber heat flux. In particular, the calorimetry outer body provides average heat flux data via circumferential channels at nine axial stations to axially resolve the heat flux distribution on the chamber outer wall. The average heat flux is found to be 5 − 12 MW/m 2 for a straight annulus and 6 − 18 MW/m 2 for a constricted annular configuration which exhibited maximum heat flux at the throat. A difference in heat flux scaling is observed near the detonation region compared to the downstream flow expansion section in the constricted geometry. By quantifying the axial distribution of heat flux, this work seeks to support the development of thermal management systems for long duration firing of detonation-based devices.

Matthew A Maybee↗
NASA NTRSDate not supplied

Average Heat Flux Measurements in a Gaseous Detonation-Based Rocket Engine

Rotating detonation rocket engines (RDREs) offer benefits over traditional deflagrative engines, including increased engine performance, compact combustion and negligible detrimental thermoacoustic instabilities due to mode-locking. Realizing the benefits of a detonation-based propulsion system will be accompanied by increased thermal loads due to a combination of heat release at elevated pressure and temperature, as well as compact heat release due to detonation. Managing these heat loads necessitates the development of thermal management strategies. As a first step towards establishing the requirements for detonation-based engine thermal management strategies, this work develops an additively manufactured (AM) water-based, axially resolved calorimeter for integration in an existing 76.8 mm outer diameter (OD) RDRE to measure average chamber heat flux. In particular, the calorimetry outer body provides average heat flux data via circumferential channels at nine axial stations to axially resolve the heat flux distribution on the chamber outer wall. The average heat flux is found to be 5 − 12 MW/m 2 for a straight annulus and 6 − 18 MW/m 2 for a constricted annular configuration which exhibited maximum heat flux at the throat. A difference in heat flux scaling is observed near the detonation region compared to the downstream flow expansion section in the constricted geometry. By quantifying the axial distribution of heat flux, this work seeks to support the development of thermal management systems for long duration firing of detonation-based devices.

Matthew A Maybee↗
NASA NTRS1982

Evaluation of advanced combustion concepts for dry NO sub x suppression with coal-derived, gaseous fuels

The emissions performance of a rich lean combustor (developed for liquid fuels) was determined for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf (7.0 to 10.3 MJ/NCM). The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher. NOx generation from NH3 is significant at ammonia concentrations significantly less than 0.5%. These levels may occur depending on fuel gas cleanup system design. CO emissions, combustion efficiency, smoke and other operational performance parameters were satisfactory. A test was completed with a catalytic combustor concept with petroleum distillate fuel. Reactor stage NOx emissions were low (1.4g NOx/kg fuel). CO emissions and combustion efficiency were satisfactory. Airflow split instabilities occurred which eventually led to test termination.

Beebe, K. W.↗
NASA NTRS1995

Modeling of microgravity combustion experiments

This program started in February 1991, and is designed to improve our understanding of basic combustion phenomena by the modeling of various configurations undergoing experimental study by others. Results through 1992 were reported in the second workshop. Work since that time has examined the following topics: Flame-balls; Intrinsic and acoustic instabilities in multiphase mixtures; Radiation effects in premixed combustion; Smouldering, both forward and reverse, as well as two dimensional smoulder.

Buckmaster, John↗
DOE OSTI2022

DNS of Hydrodynamic Instabilities of Laminar H2/O2/N2 Flames at Elevated Pressure

The atypical combustion properties of hydrogen have long provided an interest for hydrogen-enriched combustion, emphasized in recent years by increasingly stringent emission regulations. Significant work has been focused on the small hydrocarbon(C1-C3)/hydrogen blends, but much less is known about the effect of hydrogen on heavier hydrocarbon combustion, of interest for the aviation and automotive industry. In this work, we perform direct numerical simulations (DNS) of spherically expanding laminar n-dodecane/H2 flames in a constant-volume vessel at elevated pressure and a range of hydrogen seeding levels. A low Mach number model is used to represent the gas and flame dynamics allowing for a temporally varying, spatially homogeneous pressure field, while also enabling numerical evolution of the system numerically at time step sizes governed by the advective CFL limit rather than acoustic processes. We use an adaptive mesh refinement (AMR) approach to tackle the large separation of scales associated with high pressure premixed flame surfaces propagating in a laboratory-scale closed domain. We incorporate detailed chemistry and transport models for the reacting flow and focus the analysis on the effects of the pressure rise on the flame propagation characteristics and morphology as the hydrogen content is increased. The presence of fast diffusive H2 and the pressure rise results in dramatic thinning of the flame and the potential onset of hydrodyanmic and theromdiffusive instabilities. We assess the potential role of these instabilities on the ability to accurately measure laminar burning speed experimentally based on the mean flame surface propagation speed.

high pressure↗
NASA NTRS1995

Hybrid propulsion for launch vehicle boosters: A program status update

Results obtained in studying the origin and suppression of large-amplitude pressure oscillations in a 24 in. diameter hybrid motor using a liquid oxygen/hydroxylterminated polybutadiene/polycyclopentadiene propellant system are discussed. Tests conducted with liquid oxygen flow rates varying from 10 to 40 lbm/sec were designed to gauge the effectiveness of various vaporization chamber flow fields, injector designs, and levels of heat addition in suppressing high-frequency longitudinal mode oscillations. Longitudinal acoustic modes did not arise in any tests. However, initial testing revealed the presence of high-amplitude, sinusoidal, nonacoustic oscillations persisting throughout the burn durations. Analysis showed this to be analogous to chug mode instability in liquid rocket engines brought about by a coupling of motor combustion processes and the liquid oxygen feed system. Analytical models were developed and verified by test data to predict the amplitude and frequency of feed-system-coupled combustion pressure oscillations. Subsequent testing showed that increasing the feed system impedance eliminated the bulk mode instability. This paper documents the work completed to date in performance of the Hybrid Propulsion Technology for Launch Vehicle Boosters Program (NAS8-39942) sponsored by NASA's George C. Marshall Space Flight Center.

Carpenter, R. L.↗
NASA NTRS1993

Comment on tangential wave motion in unstable combustors

An attempt is made to gain insight into the injection boundary instability mechanism as captured by a simple computational tool developed for liquid rocket engines. Considerable evidence is generated supporting a hypothesis that destructive tangential-mode instabilities in liquid-fuel combustors are driven by combustion processes occurring at the propellant-wetted injector face. The driving action is found to be directly analogous to a burning solid propellant surface mechanism, responsible for solid rocket motor instabilities. Since the proposed instability mechanism has not been fundamentally confirmed, experimental research is strongly recommended.

Litchford, Ron J.↗
DOE OSTI2020

Investigation of startup, performance and cycling of a residential furnace integrated with micro-tubular flame-assisted fuel cells for micro-combined heat and power

Solid Oxide Fuel Cells (SOFCs) offer advantages for micro-Combined Heat and Power (μCHP), but currently suffer from slow startup (>1 h) and limited thermal cycling which reduces the applications, energy savings and economics. In this work, a micro-Tubular SOFC stack is integrated into a residential furnace to create a micro-Tubular Flame-assisted Fuel Cell (mT-FFC) μCHP system. A high power density of 202 mW cm -2 is reported operating in synthesis gas generated from fuel-rich combustion of natural gas/air. Unlike previous reports, instabilities in the polarization are attributed to low temperature of the oxygen reduction reaction at the cathode. The mT-FFC stack achieved peak power density in 6 min after ignition. 200 thermal cycles at an average heating rate of 215 °C.min -1 and average cooling rate of 176 °C.min -1 were conducted and a low degradation rate of 0.0325 V per 100 cycles per fuel cell was achieved. Low NO x emissions (10 ppm) and high combined efficiency is reported.

25 ENERGY STORAGE↗
NASA NTRS1983

Aerodynamic features of turbulent flames

The paper provides an interpretation for a variety of turbulent flame phenomena relating them to the aerodynamic properties of the flow field. This includes the classical notion of the breakdown of laminar flames propagation in ducts, the characteristic features of turbulent flame-fronts recorded by schlieren photography, as well as flame instabilities leading to flashback occurring in a combustion chamber behind a bluff-body flame holder. The rudimentary nature of turbulent flame propagation is shown to be associated with (1) vortex motion of a large-scale turbulent eddy, combined with (2) self-advancement of the flame at the appropriate normal burning speed and (3) the concomitant action of distributed sources of specific volume. The results furnish an insight into the macroscopic properties of the mechanism of turbulent flame propagation, rationalizing in particular its capability to sustain high flow velocities at relatively low normal burning speeds.

Oppenheim, A. K.↗
NASA NTRS2003

Two-Fluid Interface Instability Being Studied

The interface between two fluids of different density can experience instability when gravity acts normal to the surface. The relatively well known Rayleigh-Taylor (RT) instability results when the gravity is constant with a heavy fluid over a light fluid. An impulsive acceleration applied to the fluids results in the Richtmyer-Meshkov (RM) instability. The RM instability occurs regardless of the relative orientation of the heavy and light fluids. In many systems, the passing of a shock wave through the interface provides the impulsive acceleration. Both the RT and RM instabilities result in mixing at the interface. These instabilities arise in a diverse array of circumstances, including supernovas, oceans, supersonic combustion, and inertial confinement fusion (ICF). The area with the greatest current interest in RT and RM instabilities is ICF, which is an attempt to produce fusion energy for nuclear reactors from BB-sized pellets of deuterium and tritium. In the ICF experiments conducted so far, RM and RT instabilities have prevented the generation of net-positive energy. The $4 billion National Ignition Facility at Lawrence Livermore National Laboratory is being constructed to study these instabilities and to attempt to achieve net-positive yield in an ICF experiment.

Niederhaus, Charles E.↗
NASA NTRS1995

Combustion of solid fuel slabs with gaseous oxygen in a hybrid motor analog

Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated- Polybutadiene) fuel cross linked with diisocyanate was burned with GOX under various operating conditions. Large amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed line system and combustion chamber, the pressure oscillations were drastically reduced from +/- 20% of the localized mean pressure to an acceptable range of +/- 1.5%. Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations arc thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thicknesses and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented. Several tests were conducted using, simultaneously, one translucent fuel slab and one fuel slab processed with carbon black powder. The addition of carbon black did not affect the measured regression rates or surface temperatures in comparison to the translucent fuel slabs.

Chiaverini, Martin J.↗
NASA NTRS2000

Catalyst Bed Instability Within the USFE H2O2/JP-8 Rocket Engine

Orbital Sciences Corporation has been awarded a contract by NASA's Marshall Space Flight Center, in cooperation with the U.S. Air Force Research Laboratory's Military Space Plane Technology Program Office, for the Upper Stage Flight Experiment (USFE) program. Orbital is designing, developing, and will flight test a new low-cost, 10,000 lbf hydrogen peroxide/ JP-8 pressure fed liquid rocket. During combustion chamber tests at NASA Stennis Space Center (SSC) of the USFE engine, the catalyst bed showed a low frequency instability occurring as the H202 flow reached about 1/3 its design rate. This paper reviews the USFE catalyst bed and combustion chamber and its operation, then discusses the dynamics of the instability. Next the paper describes the dynamic computer model used to recreate the instability. The model was correlated to the SSC test data, and used to investigate possible solutions to the problem. The combustion chamber configuration which solved the instability is shown, and the subsequent stable operation presented.

Johnson, Curtis W.↗
NASA NTRS1995

Fuel decomposition and boundary-layer combustion processes of hybrid rocket motors

Using a high-pressure, two-dimensional hybrid motor, an experimental investigation was conducted on fundamental processes involved in hybrid rocket combustion. HTPB (Hydroxyl-terminated Polybutadiene) fuel cross-linked with diisocyanate was burned with GOX under various operating conditions. Large-amplitude pressure oscillations were encountered in earlier test runs. After identifying the source of instability and decoupling the GOX feed-line system and combustion chamber, the pressure oscillations were drastically reduced from +/-20% of the localized mean pressure to an acceptable range of +/-1.5% Embedded fine-wire thermocouples indicated that the surface temperature of the burning fuel was around 1000 K depending upon axial locations and operating conditions. Also, except near the leading-edge region, the subsurface thermal wave profiles in the upstream locations are thicker than those in the downstream locations since the solid-fuel regression rate, in general, increases with distance along the fuel slab. The recovered solid fuel slabs in the laminar portion of the boundary layer exhibited smooth surfaces, indicating the existence of a liquid melt layer on the burning fuel surface in the upstream region. After the transition section, which displayed distinct transverse striations, the surface roughness pattern became quite random and very pronounced in the downstream turbulent boundary-layer region. Both real-time X-ray radiography and ultrasonic pulse-echo techniques were used to determine the instantaneous web thickness burned and instantaneous solid-fuel regression rates over certain portions of the fuel slabs. Globally averaged and axially dependent but time-averaged regression rates were also obtained and presented.

Chiaverini, Martin J.↗