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At least 883 records · Page 49
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Effects of Wall Cooling on Supersonic Modes in High-Enthalpy Hypersonic Boundary Layers over a Cone

High-enthalpy hypersonic boundary-layer transition plays an important role in many entry/descent vehicles. Aerothermodynamic performance of these vehicles strongly depends on the transition location on the surface. However, detailed transition flow physics in these chemically reacting boundary layers are poorly understood and transition estimates during the design phase rely heavily on empirically-derived transition criteria such as 𝑹𝒆(𝜽)/𝑴(𝒆). One of the most intriguing characteristics in hypersonic boundary layers is the presence of unstable supersonic modes, first identified in the 1990s. Due to a recent surge in hypersonic applications, there has been renewed interest in studying the flow physics of supersonic modes using either theory or direct numerical simulations. This paper investigates the rise of supersonic modes in a high-enthalpy hypersonic flow over a 𝟓° half-angle cone at various wall temperatures using both quasiparallel linear stability theory and linear PSE. It was found that supersonic modes exist in all wall temperature conditions including the adiabatic wall case. Cooler wall temperature causes the second Mack mode to become an unstable supersonic mode naturally downstream of the upper-branch neutral location when the wall is sufficiently cooled. In terms of the integrated growth, the second mode is still the dominant mode. Nonetheless, supersonic modes can cause additional series of relatively weaker growth than that of the second mode beyond the peak amplitude location. According to the present linear PSE results, contrary to what was speculated in the literature, supersonic-mode pressure disturbance structure radiated into the freestream is nonacoustic in nature and the formation of unstable supersonic modes is mainly associated with the synchronization of phase speed between the instability and acoustic waves in a nonp ayer, not due to nonlinear modal interaction as suggested in the literature.

Chang, Chau-Lyan↗
NASA NTRS1991

Effect of heat release on the stability of compressible reacting mixing layer

Reacting free shear layers are of fundamental importance in many industrial systems including gas turbine combustors and rockets. Efficient propulsion systems are essential for air breathing supersonic ramjets in the high Mach number range. A limiting factor in these engines is the time for fuel and oxidizer to mix in the combustion chamber; for fast mixing, the flow must be vigorously turbulent which requires the laminar flow to be unstable. Understanding the stability characteristics of compressible reacting free shear layers is, therefore, very important and may allow one to control the flow. Low speed shear layers are highly unstable but, as chemical reaction and compressibility effects tend to stabilize them, it is important to investigate the stability of high speed reacting mixing layers. The latter consists of two fluid streams containing fuel and oxidizer respectively, and the conclusions are expected to apply, with quantitative modifications, to other shear flows, e.g., jets. Since low speed reacting cases have been studied earlier, we concentrate on the effects of Mach number and heat release. We are primarily interested in solving the stability problem over a large range of Mach number and heat release. In order to understand the effect of the heat release on the stability of this flow, one must first study the characteristics of the non-reacting flow. Inviscid theory is a reliable guide for understanding stability of compressible shear flows at moderate and large Reynolds numbers and is the basis for this work.

Shin, Dong-Shin↗
DOE OSTI2021

On the Representation of Hyporheic Exchange in Models for Reactive Transport in Stream and River Corridors

Efforts to include more detailed representations of biogeochemical processes in basin-scale water quality simulation tools face the challenge of how to tractably represent mass exchange between the flowing channels of streams and rivers and biogeochemical hotspots in the hyporheic zones. Multiscale models that use relatively coarse representations of the channel network with subgrid models for mass exchange and reactions in the hyporheic zone have started to emerge to address that challenge. Two such multiscale models are considered here, one based on a stochastic Lagrangian travel time representation of advective pumping and one on multirate diffusive exchange. The two models are formally equivalent to well-established integrodifferential representations for transport of non-reacting tracers in steady stream flow, which have been very successful in reproducing stream tracer tests. Despite that equivalence, the two models are based on very different model structures and produce significantly different results in reactive transport. In a simple denitrification example, denitrification is two to three times greater for the advection-based model because the multirate diffusive model has direct connections between the stream channel and transient storage zones and an assumption of mixing in the transient storage zones that prevent oxygen levels from dropping to the point where denitrification can progress uninhibited. By contrast, the advection-based model produces distinct redox zonation, allowing for denitrification to proceed uninhibited on part of the hyporheic flowpaths. These results demonstrate that conservative tracer tests alone are inadequate for constraining representation of mass transfer in models for reactive transport in streams and rivers.

54 ENVIRONMENTAL SCIENCES↗
NASA NTRS1992

Computational method to predict thermodynamic, transport, and flow properties for the modified Langley 8-foot high-temperature tunnel

The Langley 8 foot high temperature tunnel (8 ft HTT) is used to test components of hypersonic vehicles for aerothermal loads definition and structural component verification. The test medium of the 8 ft HTT is obtained by burning a mixture of methane and air under high pressure; the combustion products are expanded through an axisymmetric conical contoured nozzle to simulate atmospheric flight at Mach 7. This facility was modified to raise the oxygen content of the test medium to match that of air and to include Mach 4 and Mach 5 capabilities. These modifications will facilitate the testing of hypersonic air breathing propulsion systems for a wide range of flight conditions. A computational method to predict the thermodynamic, transport, and flow properties of the equilibrium chemically reacting oxygen enriched methane-air combustion products was implemented in a computer code. This code calculates the fuel, air, and oxygen mass flow rates and test section flow properties for Mach 7, 5, and 4 nozzle configurations for given combustor and mixer conditions. Salient features of the 8 ft HTT are described, and some of the predicted tunnel operational characteristics are presented in the carpet plots to assist users in preparing test plans.

Venkateswaran, S.↗
NASA NTRS1995

Fluidic Control of Nozzle Flow: Some Performance Measurements

Results are presented of an experimental program that investigated the use of a secondary air stream to control the amount of flow through a convergent-divergent nozzle. These static tests utilized high pressure, ambient temperature air that was injected at the throat of the nozzle through an annular slot. Multiple injection slot sizes and injection angles were tested. The introduction of secondary flow was made in an opposing direction to the primary flow and the resulting flow field caused the primary stream to react as though the physical throat size had been reduced. The percentage reduction in primary flow rate was generally about twice the injected flow rate. The most effective throttling was achieved by injecting through the smallest slot in an orientation most nearly opposed to the approaching primary flow. Thrust edliciency, as measured by changes in nozzle thrust coefficient, was highest at high nozzle pressure ratios, NPR. The static test results agreed with predictions obtained prior from PABSD, a fully viscous computational fluid dynamics program. Since use of such an injection system on gas turbine engine exhaust nozzles would be primarily at high NPRs, it was concluded that fluidic control holds promise for reducing nozzle weight and complexity on future systems.

Federspiel, John↗
DOE OSTI2021

Partial Premixing Effects on the Reacting Jet of a High-Pressure Axially Staged Combustor

The effects of partial premixing on a reacting jet-in-crossflow is investigated in a five atmosphere axially staged combustor at stationary gas turbine relevant conditions. The facility consists of a dump style headend burner that provides a crossflow with a quasi-uniform velocity and temperature profile to the axial stage to isolate the effects of the jet-in-crossflow. The headend burner is run with methane and air at a lean equivalence ratio to match industry emission standards. For this work, the total air to the headend and axial stage is kept constant, and fuel is split between the headend and axial stage to represent different gas turbine loading conditions. For the cases analyzed, the fuel split to the axial stage went up to 25%. The axial stage consists of an optically accessible test section with a coaxial injector that provides variability to how long the methane and air can mix before entering the facility. Three different premixed levels are studied: fully premixed, nonpremixed, and partially premixed. The flow-field characteristics of the reacting jet-in-crossflow are analyzed using particle image velocimetry (PIV), and flame behavior is quantified by employing CH* chemiluminescence. NO measurements are made at the exit of the facility using a Horiba emissions analyzer. Furthermore, two different flames are observed: flames that burn in the leeward recirculation region and flames that burn at the core of the jet.

03 NATURAL GAS↗
NASA NTRS1974

Multiple eigenvalues of sound-absorbing circular and annular ducts

Eigenvalue equations of the form D(lambda) = 0 are studied for acoustic modes in circular and annular ducts without flow. The ducts have locally reacting walls with arbitrary wall admittance. It is shown that circular ducts may have double eigenvalues, or wall admittances where both D(lambda) and D prime (lambda) are zero. These double eigenvalues are formed by the coalescence of the lowest-order eigenfunction with some higher-order eigenfunction. The eigenfunctions, or acoustic modes, associated with the double eigenvalues are found to be orthogonal to themselves so that the standard expansion formula for the acoustic field, which is obtained from a separation-of-variables analysis, is invalid. An alternate expansion formula for the acoustic field is derived. This alternate formula shows that the coalesced acoustic mode is linearly amplified by the transmission distance and exponentially attenuated at the expected rate. The expansion formula also reveals a new eigenfunction, or acoustic mode, which is exponentially attenuated at the same rate as the coalesced eigenfunction.

Zorumski, W. E.↗
NASA NTRS1993

Modifications Of A Commercial Spray Gun

Commercial spray gun modified to increase spray rate and make sprayed coats more nearly uniform. Consists of gun head and pneumatic actuator. Actuator opens valves for two chemical components, called "A" and "B," that react to produce foam. Components flow through orifices, into mixing chamber in head. Mixture then flows through control orifice to spray tip. New spray tip tapered to reduce area available for accumulation of foam and makes tip easier to clean.

Allen, Peter B.↗
NASA NTRS2004

Validation of an Impedance Education Method in Flow

This paper reports results of a research effort to validate a method for educing the normal incidence impedance of a locally reacting liner, located in a grazing incidence, nonprogressive acoustic wave environment with flow. The results presented in this paper test the ability of the method to reproduce the measured normal incidence impedance of a solid steel plate and two soft test liners in a uniform flow. The test liners are known to be locally react- ing and exhibit no measurable amplitude-dependent impedance nonlinearities or flow effects. Baseline impedance spectra for these liners were therefore established from measurements in a conventional normal incidence impedance tube. A key feature of the method is the expansion of the unknown impedance function as a piecewise continuous polynomial with undetermined coefficients. Stewart's adaptation of the Davidon-Fletcher-Powell optimization algorithm is used to educe the normal incidence impedance at each Mach number by optimizing an objective function. The method is shown to reproduce the measured normal incidence impedance spectrum for each of the test liners, thus validating its usefulness for determining the normal incidence impedance of test liners for a broad range of source frequencies and flow Mach numbers. Nomenclature

Watson, Willie R.↗
NASA NTRS1994

A random distribution reacting mixing layer model

A methodology for simulation of molecular mixing, and the resulting velocity and temperature fields has been developed. The ideas are applied to the flow conditions present in the NASA Lewis Research Center Planar Reacting Shear Layer (PRSL) facility, and results compared to experimental data. A gaussian transverse turbulent velocity distribution is used in conjunction with a linearly increasing time scale to describe the mixing of different regions of the flow. Equilibrium reaction calculations are then performed on the mix to arrive at a new species composition and temperature. Velocities are determined through summation of momentum contributions. The analysis indicates a combustion efficiency of the order of 80 percent for the reacting mixing layer, and a turbulent Schmidt number of 2/3. The success of the model is attributed to the simulation of large-scale transport of fluid. The favorable comparison shows that a relatively quick and simple PC calculation is capable of simulating the basic flow structure in the reacting and nonreacting shear layer present in the facility given basic assumptions about turbulence properties.

Jones, Richard A.↗
NASA NTRS1992

Planar laser-induced fluorescence imaging of OH in the exhaust of a bi-propellant thruster

Planar laser-induced fluorescence imaging of the hydroxyl radical has been performed on the flow produced by the exhaust of a subscale H2/O2 fueled bi-propellant rocket engine. Measurements were made to test the feasibility of OH (0,0) and (3,0) excitation strategies by using injection seeded XeCl and KrF excimer lasers, respectively. The flow is produced with hydrogen and oxygen reacting at a combustor chamber pressure of 5 atm which then exhausts to the ambient. The hydroxyl concentration in the exhaust flow is approximately 8 percent. Fluorescence images obtained by pumping the Q1(3) transition in the (0,0) band exhibited very high signals but also showed the effect of laser beam absorption. To obtain images when pumping the P1(8) transition in the (3,0) band it was necessary to use exceptionally fast imaging optics and unacceptably high intensifier gains. The result was single-shot images which displayed a signal-to-noise ratio of order unity or less when measured on a per pixel basis.

Paul, Phillip H.↗
NASA NTRS1992

A numerical investigation of hydrogen combustion in Mach 2 flow

A numerical study is conducted of a reacting flowfield generated by a single hydrogen jet injected transversely behind a rearward facing step. A nonreacting mixing study was followed by a reacting simulation. The computation simulates a current experimental condition. The validation of the numerical solution was limited to a comparison of the wall-pressure distribution along the wall. The comparison indicates agreement between the experiment and the prediction with a maximum error of 7 percent. The addition of the reaction responsible for HO2 production (H + O2 + M = HO2 + M) to an initially simplified chemical mechanism reduced the exponential growth of the free radicals and reduced the amount of heat released by combustion; as a result, the flowfield was significantly changed. A reversed flow was obtained in limited regions of the flow.

Segal, Corin↗
NASA NTRS2019

Physical Simulation of Rocket Exhaust Aerodynamics Using Heated Ethane: Conceptual Foundations

Heated ethane (C2H6) has been proposed as an alternative to inert gases for use as a motive fluid in the experimental simulation of rocket exhaust plumes. By adjusting stagnation temperature, the isentropic exponent of ethane can be tuned to approximate those produced by common rocket propellants including hydrogen, hypergols, alcohols, and hydrocarbons. As a result, ethane can be made to follow a nozzle expansion process which is nearly identical to realistic rocket engine flow fields. Additionally, its high auto-ignition temperature and resistance to condensation enable the testing of expansion ratios much larger than conventional inertgas testing. NASA SSC has performed quasi-one-dimensional analyses using the Chemical Equilibrium with Applications (CEA) code as a preliminary means to compare flow fields produced by non-reacting ethane to those of reacting combustion products. A LO2/LH2 rocket engine operating at a chamber pressure of 5.0 MPa and a mixture ratio of 6.1 was used as an example case to demonstrate ethane’s efficacy as a simulant. Errors for key similarity parameters were compared to legacy cold-flow test methods. Additional errors induced by machining tolerances and chemical impurities were also examined. Results suggest that at a 3% geometric scale and ~500 K ethane stagnation temperature, an error of less than 2.5% throughout the flow field is realistically achievable along the dimensions of Mach number, Reynolds number, pressure ratio, and isentropic exponent. The development of an experimental test bed for validation of this configuration is currently underway.

Jones, Daniel R.↗
NASA NTRS1993

Nonequilibrium effects on the aerothermodynamics of transatmospheric and aerobraking vehicles

A 3D CFD algorithm is used to study the effect of thermal and chemical nonequilibrium on slender and blunt body aerothermodynamics. Both perfect gas and reacting gas air models are used to compute the flow over a generic transatmospheric vehicle and a proposed lunar transfer vehicle. The reacting air is characterized by a translational-rotational temperature and a vibrational-electron-electronic temperature and includes eight chemical species. The effects of chemical reaction, vibrational excitation, and ionization on lift-to-drag ratio and trim angle are investigated. Results for the NASA Ames All-body Configuration show a significant difference in center of gravity location for a reacting gas flight case when compared to a perfect gas wind tunnel case at the same Mach number, Reynolds number, and angle of attack. For the same center of gravity location, the wind tunnel model trims at lower angle of attack than the full-scale flight case. Nonionized and ionized results for a proposed lunar transfer vehicle compare well to computational results obtained from a previously validated reacting gas algorithm. Under the conditions investigated, effects of weak ionization on the heat transfer and aerodynamic coefficients were minimal.

Hassan, Basil↗
NASA NTRSDate not supplied

A Computational Study of Plume Modeling for Space Launch System Abort Scenarios

A series of viscous CFD simulations depicting Space Launch System (SLS) post Mode-1abort scenarios are conducted and analyzed. The primary purpose of these simulations is to study the effect of launch abort vehicle attitude control motor (ACM) and abort motor(AM) plume modeling fidelity on the drag of the aborted core stage. The simulations are first conducted with a fully-coupled, multi-species and chemically reacting model for the freestream gases and the solid rocket abort motor combustion products. Following this, equivalent species are created for the ACM and AM exhaust gases. These equivalent species aim to provide near the same results for core stage drag as the chemically reacting model at a significantly lower computational cost. Finally, the exhaust gases are modeled as calorically perfect air. It was concluded that under most scenarios, both the equivalent species and the perfect air models predict the drag of the aborted core stage sufficiently close to that predicted by the chemically reacting simulations. In these cases, the flow and geometry conditions did not induce significant changes in the composition of the ‘true’ exhaust gas via afterburning of the combustion products. For cases where this was not the case, the properties of the essentially frozen equivalent species and perfect air models deviated from the true mixture, yielding unsatisfactory results. Finally, the accuracy improvements by using the chemically reacting model are weighed against the significant increase in associated computational costs. It was found that the simplified exhaust gas models are at least 4 times less expensive than the chemically reacting model. This may justify the judicious use of these simple models in future work.

CFD↗
NASA NTRSDate not supplied

A Computational Study of Plume Modeling for Space Launch System Abort Scenarios

A series of viscous CFD simulations depicting Space Launch System (SLS) post Mode-1abort scenarios are conducted and analyzed. The primary purpose of these simulations is to study the effect of launch abort vehicle attitude control motor (ACM) and abort motor(AM) plume modeling fidelity on the drag of the aborted core stage. The simulations are first conducted with a fully-coupled, multi-species and chemically reacting model for the freestream gases and the solid rocket abort motor combustion products. Following this, equivalent species are created for the ACM and AM exhaust gases. These equivalent species aim to provide near the same results for core stage drag as the chemically reacting model at a significantly lower computational cost. Finally, the exhaust gases are modeled as calorically perfect air. It was concluded that under most scenarios, both the equivalent species and the perfect air models predict the drag of the aborted core stage sufficiently close to that predicted by the chemically reacting simulations. In these cases, the flow and geometry conditions did not induce significant changes in the composition of the ‘true’ exhaust gas via afterburning of the combustion products. For cases where this was not the case, the properties of the essentially frozen equivalent species and perfect air models deviated from the true mixture, yielding unsatisfactory results. Finally, the accuracy improvements by using the chemically reacting model are weighed against the significant increase in associated computational costs. It was found that the simplified exhaust gas models are at least 4 times less expensive than the chemically reacting model. This may justify the judicious use of these simple models in future work.

CFD↗
NASA NTRS1995

Investigation Flow Uniformity in a Supersonic Duct with High Enthalpy Flows

Flow uniformity in a high enthalpy facility is investigated. The scramjet research facility is composed of a rectangular combustor duct connected to a 100 MW electric arc air heater. The Mach 3.3 flow is accelerated through a two-dimensional contoured nozzle. Instream measurements were made with water-cooled Pitot probes and stagnation point heat flux gages at stream enthalpy levels ranging from 4 to 7 Mj/kg. Flow surveys were made on the flow centerline and off centerline in order to measure the three dimensional uniformity of the flow in the rectangular duct. Measurements indicated that although the flow in the aspect ratio 6:1 duct was relatively uniform on the centerline, three dimensional viscous effects were apparent near the corners. Flow through the nozzle and constant area duct was modeled computationally using a two dimensional, Navier-Stokes, reacting gas code. The computations predict that the flow in the test section is in vibrational equilibrium. The computed and measured Pitot pressure and heat flux profiles are in reasonable agreement with the experimental data.

Balboni, John↗
NASA NTRS1991

Tuned optimization of extended reacting acoustic liners

A finite element Galerkin formulation was employed to study the optimum attenuation and reflection characteristics of acoustic waves propagating in tw0 dimensional straight ducts with extended reacting absorbing walls without a mean flow. The reflection and transmission of acoustic energy at the entrance and the exit of the duct were determined by coupling the finite element solutions in the absorbing portion of the duct to the eigenfunctions of an infinite, uniform, hard wall duct. In the frequency range where the duct height and acoustic wave length are nearly equal, power attenuation contours were examined to determine conditions for minimizing acoustic transmission through the duct. The extended reacting liners were found to significantly minimize the reflective characteristics of the duct rather than to increase absorption in the liner. Thus, extended reacting wall liner properties can be theoretically chosen to yield large impedance mismatches in an open straight duct.

Baumeister, Kenneth J.↗