Monograph records
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A Caltech Library Repository Feedhttp://www.rssboard.org/rss-specificationpython-feedgenenThu, 30 Nov 2023 18:13:52 +0000Effect of NACA Injection Impeller on Mixture Distribution of Double-Row Radial Aircraft Engine
https://resolver.caltech.edu/CaltechAUTHORS:20140805-132115205
Authors: Marble, Frank E.; Ritter, William K.; Miller, Mahlon A.
Year: 1945
The NACA injection impeller was developed to improve the mixture distribution of aircraft engines by discharging the fuel from a centrifugal supercharger impeller and thus to promote a thorough mixing of fuel and charge air. Experiments with a double-row radial aircraft engine indicated that for the normal range of engine power the NACA injection impeller provided marked improvement in mixture distribution over the standard spray-bar injection system used in the same engine. The mixture distribution at cruising conditions was excellent; at 1200, 1500, and 1700 brake horsepower, the differences between the fuel-air ratios of the richest and the leanest cylinders were reduced to approximately one-third their former values.https://authors.library.caltech.edu/records/ctqnj-xnw02Analysis of cooling limitations and effect of engine-cooling improvements on level-flight cruising performance of four-engine heavy bomber
https://resolver.caltech.edu/CaltechAUTHORS:MARnacarpt860
Authors: Marble, Frank E.; Miller, Mahlon E.; Bell, E. Barton
Year: 1946
The NACA has developed means, including an injection impeller and ducted head baffles, to improve the cooling characteristics of the 33500-cubic-inch-displacement radial engines installed in a four-engine heavy bomber. The improvements afforded proper cooling of the rear-row exhaust-valve seats for a wide range of cowl-flap angles, mixture strength, and airplane speeds. The results of flight tests with this airplane are used as a basis for a study to determine the manner and the extent to which the airplane performance was limited by engine cooling. By means of this analysis for both the standard airplane and the airplane with engine-cooling modifications, comparison of the specific range at particular conditions and comparison of the cruising-performance limitations were made.
The analysis of level-flight cruising performance of the airplane with both the standard- and the modified-engine installations indicated that the maximum cruising economy is attained at the minimum brake specific fuel consumption when engine cooling under these conditions is possible. Operation at lean mixtures, high altitudes, and large gross weights was limited for the standard airplane by engine cooling at the point where larger cowl-gap openings increase the power required for level flight at such a rate that the additional cooling air available is insufficient to cool the engine when developing the additional power. When cooling becomes impossible at the minimum brake specific fuel consumption, the maximum cruising economy is obtained with a cowl-flap angle of approximately 6[degrees] and with the leanest mixture (above the stoichiometric value) giving satisfactory
engine cooling.
Comparison of the calculated perfomance of the standard and the modified airplane indicated that cooling improvements increased the maximum specfic range as much as 38 percent
for operation where wide cowl-flap angles and enriched mixtures are required to cool the standard airplane. Corresponding increases in cruising range were calculated for flights in which conditions allowing large increases in cruising economy were encountered. The cooling improvements allow either an increase of more than 10,000 feet in operating altitude at a given airplane weight or a gross-weight increase of from 10,000 pounds at sea lerel to 35,000 pounds at all operating altitudes above 10,000 feet.https://authors.library.caltech.edu/records/d1cqm-kj337Effect of the NACA Injection Impeller on the Mixture Distribution of a Double-Row Radial Aircraft Engine
https://resolver.caltech.edu/CaltechAUTHORS:MARnacatn1069
Authors: Marble, Frank E.; Ritter, William K.; Miller, Mahlon A.
Year: 1946
The NACA injection impeller was developed to improve the mixture distribution of aircraft engines by discharging the fuel from a centrifugal supercharger impeller, thus promoting a thorough mixing of fuel and charge air. Tests with a double-row radial aircraft engine indicated that for the normal range of engine power the NACA injection impeller provided marked improvement in mixture distribution over the standard spray-bar injection system used in the same engine. The mixture distribution at cruising conditions was excellent; at 1200, 15OO, and 1700 brake horsepower, the differences between the fuel-air ratios of the richest and the leanest cylinders were reduced to approximately one-third their former values. The maximum cylinder temperatures were reduced about 30 [degrees] F and the temperature distribution was improved by approximately the
degree expected from the improvement in mixture distribution. Because the mixture distribution of the engine tested improves slightly at engine powers exceeding 1500 brake horsepower and because the effectiveness of the particular impeller diminished slightly at high rates of fuel flow, the improvement in mixture distribution at
rated power and rich mixtures was less than that for other conditions.
The difference between the fuel-air ratios of the richest and the leanest cylinders of the engine using the standard spray bar was so great that the fuel-air ratios of several cylinders were well below the chemically correct mixture, whereas other cylinders were operating at rich mixtures. Consequently, enrichment to improve engine cooling actually increascd some of the critical temperatures. The uniform mixture distribution providod by the injection impeller restored the normal response of cylinder temperatures to mixture enrichnent.https://authors.library.caltech.edu/records/jnvkf-b4h33Analytical Investigation of Some Three-Dimensional Flow Problems in Turbomachines
https://resolver.caltech.edu/CaltechAUTHORS:MARnacatn2614
Authors: Marble, Frank E.; Michelson, Irving
Year: 1952
One problem encountered in the theory of turbomachines is that of calculating the fluid velocity components when the inner and outer boundaries of the machine as well as the shape of or forces imparted by the blade row are given. The present paper discusses this problem under the restrictions that the fluid is inviscid and incompressible and that the blade rows consist of an infinite number of infinitely thin blades so that axially symmetric flow is assumed.
It is shown, in general, that the velocity components in a plane through the turbomachine axis may be expressed in terms of the angular momentum and the leading-edge blade force normal to the stream surfaces. The relation is a nonlinear differential equation to which analytic solutions may be obtained conveniently only after the introduction of linearizing assumptions. A quite accurate linearization is effected through assuming an approximate shape of the stream surfaces in certain nonlinear terms.
The complete linearized solution for the axial turbomachine is given in such form that blade loading, blade shape, distribution of angular momentum, or distribution of total head may be prescribed. Calculations for single blade rows of aspect ratio 2 and 2/3 are given for a radius ratio of 0.6. They indicate that the process of formation of the axial velocity profile may extend both upstream and downstream of a high-aspect-ratio blade row, while for low aspect ratios the major portion of the three-dimensional flow occurs within the blade row itself. When the through-flow velocity varies greatly from its mean value, the simple linearized solution does not describe the flow process adequately and a more accurate solution applicable to such conditions is suggested.
The structure of the first-order linearized solution for the axial turbomachine suggested a further approximation employing a minimizing operation. The simplicity of this solution permits the discussion of three interesting problems: Mutual interference of neighboring blade rows in a multistage axial turbomachine, solution for a single blade row of given blade shape, and the solution for this blade row operating at a condition different from the design condition.
It is found that the interference of adjacent blade rows in the multistage turbomachine may be neglected when the ratio of blade length to the distance between centers of successive blade rows is 1.0 or less. For values of this ratio in excess of 3.0, the interference may be an important influence. The solution for the single blade row indicated that, for the blade shape considered, the distortion of the axial velocity profile caused by off-design operation is appreciably less for low- than for high-aspect-ratio blades.
To obtain some results for a mixed-flow turbomachine comparable with those for the axial turbomachine as well as to indicate the essential versatility of the method of linearizing the general equations, completely analogous theoretical treatment is given for a turbomachine whose inner and outer walls are concentric cones with common apex and whose flow is that of a three-dimensional source or sink. A particular example for a single rotating blade row is discussed where the angular momentum is prescribed similarly to that used in the examples for the axial turbomachine.https://authors.library.caltech.edu/records/zt6sp-jyw37Unsteady flows in axial turbomachines
https://resolver.caltech.edu/CaltechAUTHORS:20110204-111252490
Authors: Rannie, W. Duncan; Marble, Frank E.
Year: 1957
Of the various unsteady flows that occur in axial turbomachines certain asymmetric disturbances, of wave length large in comparison with blade spacing, have become understood to a certain extent. These disturbances divide themselves into two categories: self-induced oscillations and forced disturbances. A special type of propagating stall appears as a self-induced disturbance; an asymmetric velocity profile introduced at the compressor inlet
constitutes a forced disturbance.
Both phenomena have been treated from a unified theoretical point of view in which the asymmetric disturbances are linearized and the blade characteristics are assumed quasi-steady. Experimental results are in essential agreement with this theory wherever the limitations of the theory are satisfied. For the self-induced disturbances and the more interesting examples of the forced disturbances, the dominant blade characteristic is the dependence of total pressure loss, rather than the turning angle, upon the local blade inlet angle.https://authors.library.caltech.edu/records/xemjt-3p211The role of approximate analytical results in the study of two-phase flow in nozzle
https://resolver.caltech.edu/CaltechAUTHORS:20101222-101345306
Authors: Marble, Frank E.
Year: 1967
The small slip approxitnation to the theory of two-phase flow in rocket nozzles is reviewed to show that the inaccuracies associated with drag and heat transfer laws, and those associated with the fundamental approximation,
are independent and that the former may be removed algebraic1y. Selected applications ofthe approximate theory are discussed to indicate that these stress the nature of the dependence of the results upon the relevant physical parameters and the possible consequence of scaling
laws, rather than numerical accuracy too often limited by inaccurate initial data. It is suggested that approximate analytical results may offer much more assistance to the rocket engineer than has yet been used to advantage.https://authors.library.caltech.edu/records/gz0k7-20g34Large building fires - experiments and analysis
https://resolver.caltech.edu/CaltechAUTHORS:20110203-131719678
Authors: Zukoski, Edward E.; Marble, Frank E.; Ranie, W. Duncan
Year: 1970
Because of its inherent complexity and detail, as well as its rather tenuous relationship to existing combustion theory, the propagation of uncontrolled fires in large buildings remains one of the unsolved problems facing our cities. On October 13, 1969 (see Appendix), a fire in a
Los Angeles apartment claimed the lives of eight people and sent more than a score to the hospital for various degrees of burn and smoke inhalation. As the fire developed, flames spread quickly up the main stairwell, blocking exits from apartment units, forcing some to jump from upper floors. Within a matter of minutes, all three floors were so involved in fire that normal escape was impossible.
Our lack of quantitative knowledge about the propagation of building fire has a more widespread effect than such disasters. It is a major factor in preserving archaic and inappropriate building codes; it places a severe
limit on architectural innovation because fire hazards in novel structures cannot be evaluated quantitatively. This is a truly serious restriction in an era where low-cost multiple dwellings are in urgent need.https://authors.library.caltech.edu/records/qa99p-1t908A theory of two-dimensional airfoils with strong inlet flow on the upper surface
https://resolver.caltech.edu/CaltechAUTHORS:20110208-150155345
Authors: Serdengecti, Sedat; Marble, Frank E.
Year: 1970
The two-dimensional theory of airfoils with arbitrarily strong inlet
flow into the upper surface was examined with the aim of developing a thin-airfoil
theory which is valid for this condition. Such a theory has, in fact,
been developed and reduces uniformly to the conventional thin-wing theory
when the inlet flow vanishes. The integrals associated with the arbitrary
shape, corresponding to the familiar Munk integrals, are somewhat more
complex but not so as to make calculations difficult.
To examine the limit for very high ratios of inlet to free-stream
velocity, the theory of the Joukowski airfoil was extended to incorporate
an arbitrary inlet on the upper surface. Because this calculation is exact,
phenomena observed in the limit cannot be attributed to the linearized calculation. These results showed that airfoil theory, in the conventional sense,
breaks down at very large ratios of inlet to free-stream velocity. This
occurs where the strong induced field of the inlet dominates the free-stream
flow so overwhelmingly that the flow no longer leaves the trailing edge but
flows toward it. Then the trailing edge becomes, in fact a leading edge
and the Kutta condition is physically inapplicable. For the example in this
work, this breakdown occurred at a ratio of inlet to free-stream velocity
of about 10. This phenomena suggests that for ratios in excess of the
critical value, the flow separates from the trailing edge and the circulation
is dominated by conditions at the edges of the inlet.https://authors.library.caltech.edu/records/a5gkn-s1p04The generation of noise by the fluctuations in gas temperature into a turbine
https://resolver.caltech.edu/CaltechAUTHORS:20110126-092110939
Authors: Cumpsty, Nicholas A.; Marble, F. E.
Year: 1974
An actuator disc analysis is used to calculate the pressure
fluctuations produced by the convection of temperature
fluctuations (entropy waves) into one or more rows of blades.
The perturbations in pressure and temperature must be small,
but the mean flow deflection and acceleration are generally
large. The calculations indicate that the small temperature
fluctuations produced by combustion chambers are sufficient
to produce large amounts of acoustic power.
Although designed primarily to calculate the effect of
entropy waves, the method is more general and is able to
predict the pressure and vorticity waves generated by
upstream or downstream going pressure waves or by vorticity
waves impinging on blade rows.https://authors.library.caltech.edu/records/mhvb2-km066Response of a nozzle to an entropy disturbance example of thermodynamically unsteady aerodynamics
https://resolver.caltech.edu/CaltechAUTHORS:20110208-105530510
Authors: Marble, Frank E.
Year: 1975
The larger number of problems that qualify as unsteady aerodynamics
relate to non-uniform motion of surfaces -- such as pitching of
airfoils -- or the correspondingly non-uniform motion of a fluid about a
surface -- such as a gust passing over an airfoil. Experiment and analysis
concerning these problems aims to determine the non-steady forces
or surface stresses on the object. These may be thought of as "kinematically" non-steady problems. Another class of problems presents itself
when the undisturbed gas stream temperature (or density) is non-steady
although the velocity and pressure are steady; such non-uniformities are
associated with entropy variations from point to point of the stream. In a
locally adiabatic flow these entropy variations are transported with the
stream, and when a fixed boundary -- such as an airfoil -- is encountered,
the flow field undergoes a non-steady change because the density variations
alter the pressure field -- or the stresses at the boundaries. This happens
in spite of the fact that the undisturbed free -stream velocity field and the
surface boundaries of the flow are independent of time. A gas turbine blade, for example, will experience a time-dependent load simply because
of temperature fluctuations in the combustion products flowing over it, although
the angle of attack is independent of time. We shall call these
"thermodynamically" unsteady flows in contrast with the more familiar
kinematically unsteady flows.https://authors.library.caltech.edu/records/my18d-yqk54The Coherent Flame Model for Turbulent Chemical Reactions
https://resolver.caltech.edu/CaltechAUTHORS:20101210-105056861
Authors: Marble, Frank E.; Broadwell, James E.
Year: 1977
A description of the turbulent diffusion flame is proposed in which the flame structure is composed of a distribution of laminar diffusion flame elements, whose thickness is small in comparison with the large eddies. These elements retain their identity during the flame development; they are strained in their own plane by the gas motion, a process that not only extends their surface area, but also establishes the rate at which a flame element consumes the reactants. Where this flame stretching process has produced a high flame surface density, the flame area per unit volume, adjacent flame elements may consume the intervening reactant, thereby annihilating both flame segments. This is the flame shortening mechanism which, in balance with the flame stretching process, establishes the local level of the flame density. The consumption rate of reactant is then given simply by the product of the local flame density and the reactang consumption rate per unit area of flame surface. The proposed description permits a rather complete separation of the turbulent flow structure, on one hand, and the flame structure, on the other, and in this manner permits the treatment of reactions with complex chemistry with a minimum of added labor. The structure of the strained laminar diffusion flame may be determined by analysis, numerical computation, and by experiment without significant change to the model.https://authors.library.caltech.edu/records/k89f0-pfd09Theoretical Analysis of Nitric Oxide Production in a Methane/Air Turbulent Diffusion Flame
https://resolver.caltech.edu/CaltechAUTHORS:20101209-090337380
Authors: Marble, Frank E.
Year: 1980
The coherent flame model is applied to the methane-air turbulent diffusion flame with the objective of describing the production of nitric oxide. The example of a circular jet of methane discharging into a stationary air atmosphere is used to illustrate application of the model. In the model, the chemical reactions take place in laminar flame elements which are lengthened by the turbulent fluid motion and shortened when adjacent flame segments consume intervening reactant. The rates with which methane and
air are consumed and nitric oxide generated in the strained laminar flame are computed numerically in an independent calculation. The model predicts nitric oxide levels of approximately 80 parts per million at the end of the flame generated by a 30.5 cm (1 foot) diameter jet of methane issuing at 3.05 x 10^3 cm/sec (100 ft/sec). The model also
predicts that this level varies directly with the fuel jet diameter and inversely with the jet velocity. A possibly important nitric oxide production mechanism, neglected in
the present analysis, can be treated in a proposed extension to the model.https://authors.library.caltech.edu/records/rmgmj-djj95Experiments concerning the mechanism of flame blowoff from bluff bodies
https://resolver.caltech.edu/CaltechAUTHORS:20110203-125953778
Authors: Zukoski, Edward E.; Marble, Frank E.
Year: 1983
The general problem of flame stabilization on bluff objects centers about the determination of the maximum stream velocity at which stable combustion may be achieved for various flame holder geometries, gas mixtures and conditions of the approaching combustible stream. Since the process involves both gas dynamic problems and chemical kinetic problems of great complexity, the most reasonable approach is one of similarity, that is, to determine under what conditions the behavior of one flame holder is similar to the behavior of another one. Because a very large number of physical and chemical variables is involved in a combustion problem, similarity conditions can be formulated most easily after experimental investigations have indicated which parameters or groups exert little influence on the mechanism and hence may be neglected. The experiments
described in this paper were conducted with the object
of clarifying the role of the more important parameters
in the flame holding mechanism. The results indicate
that a relatively simple formulation of the similarity conditions may be obtained in which the fluid mechanical
parameters and chemical parameters are effectively separated.https://authors.library.caltech.edu/records/an2gy-r1y81Progress toward shock enhancement of supersonic combustion processes
https://resolver.caltech.edu/CaltechAUTHORS:20101201-110408219
Authors: Marble, Frank E.; Hendricks, Gavin J.; Zukoski, Edward E.
Year: 1987
In air breathing propulsion systems for flight at Mach numbers 7 to 20, it is generally accepted that the combustion processes will be carried out at supersonic velocities with respect to the engine. The resulting brief residence time places a premium on rapid mixing of the
fuel and air. To address this issue we &re investigating a mechanism for enhancing the rate of mixing between air and hydrogen fuel over rates that are expected in shear layers and jets. The mechanism rests upon the strong vorticity induced at the interface between a light and heavy gas by an intense pressure gradient. The specific phenomenon under investigation is the rapid mixing induced by interaction of a weak oblique shock with a cylindrical jet of hydrogen embedded in air. The status of our investigations is described in three parts: a) shock tube investigation of the distortion and mixing induced by shock waves
impinging on cylindric of hydrogen embedded in air, b) the molecular mixing and chemical reaction in large vortices, periodically formed in a channel, and c) two-dimensional non-steady and three-dimensional steady numerical studies of shock interaction with cylindrical volumes of hydrogen in air.https://authors.library.caltech.edu/records/q52jm-t0410An Experimental and Numerical Investigation of Swirling Flows in a Rectangular Nozzle
https://resolver.caltech.edu/CaltechAUTHORS:20101203-145603258
Authors: Sobota, Thomas H.; Marble, Frank E.
Year: 1987
The high thrust to weight ratios now possible for military aircraft have made thrust vector pitch control more attractive and versatile than aerodynamic surface pitch control. Use of a rectangular nozzle is a natural consequence because articulation and sealing problems are less formidable than for conventional circular ones. The rectangular nozzle offers the additional possibility that the exhaust may mix rapidly with the ambient air and thereby reduce the radiative signature of the exhaust. A detailed experimental investigation is described, which demonstrates that the formation of axial vortices in the nozzle is dependent on the vorticity distribution at the turbine exhaust. Further, three mechanisms which provide for the formation of axial vortices are identified. A parallel computational investigation was carried out which not only confirmed the relationship between the turbine exhaust vorticity and the vortex pattern formed in the nozzle but also provided details of the flow field between the turbine discharge and the nozzle exit. On the basis of this more detailed understanding, it is now possible to tailor the vortex distribution at the nozzle exit by design of the turbine discharge and the intervening passage.https://authors.library.caltech.edu/records/y7w3d-08a25Shock enhancement and control of hypersonic mixing and combustion
https://resolver.caltech.edu/CaltechAUTHORS:20101209-134118457
Authors: Marble, F. E.; Zukoski, E. E.; Jacobs, J. W.; Hendricks, G. J.; Waitz, I. A.
Year: 1990
The possibility that shock enhanced mixing can
substantially increase the rate of mixing between
coflowing streams of hydrogen and air has been
studied in experimental and computational investigations.
Early numerical computations indicated that
the steady interaction between a weak shock in air
with a coflowing hydrogen jet can be well approximated
by the two-dimensional time-dependent interaction
between a weak shock and an initially circular
region filled with hydrogen imbedded in air. An experimental
investigation of the latter process has been
carned out in the Caltech 17 Inch Shock Tube in experiments
in which the laser induced fluorescence of
byacetyl dye is used as a tracer for the motion of the
helium gas after shock waves have passed across the
helium cylinder. The flow field has also been studied
using an Euler code computation of the flow field.
Both investigations show that the shock impinging
process causes the light gas cylinder to split into two
parts. One of these mixes rapidly with air and the
other forms a stably stratified vortex pair which mixes
more slowly; about 60% of the light gas mixes rapidly
with the ambient fluid. The geometry of the flow field
and the mixing process and scaling parameters are
discussed here. The success of this program encouraged
the exploration of a low drag injection system in
which the basic concept of shock generated streamwise
vorticity could be incorporated in an injector for
a Scramjet combustor at Mach numbers between 5
and 8. The results of a substantial computational
program and a description of the wind tunnel model and preliminary experimental results obtained in the
High Reynolds Number Mach 6 Tunnel at NASA Langley
Research Center are given here.https://authors.library.caltech.edu/records/4b59h-1v374An investigation of a contoured wall injector for hypervelocity mixing augmentation
https://resolver.caltech.edu/CaltechAUTHORS:20101130-101649129
Authors: Waitz, Ian-A; Marble, Frank E.; Zukoski, Edward E.
Year: 1991
An experimental and computational investigation of a
contoured wall fuel injector is presented. The injector
was aimed at enabling shock-enhanced mixing for the
supersonic combustion ramjet engines currently envisioned for applications on hypersonic vehicles. Three-dimensional flow field surveys, and temporally resolved planar Rayleigh scattering measurements are presented for Mach 1.7 helium injection into Mach 6 air. These experimental data are compared directly with a three-dimensional Navier-Stokes simulation of the flow about the injector array. Two dominant axial vorticity sources are identified and characterized. The axial vorticity produced strong convective mixing of the injectant with the freestream. Shock-impingement was particularly effective as it assured seeding of baroclinic vorticity directly on the helium/air interface. The vorticity coalesced into a counter-rotating vortex pair of a sense which produced migration of the helium away from the wall. The influences of spatial averaging on the representation of the flow field as well as the importance of the fluctuating component of the flow in producing molecularly-mixed fluid are addressed.https://authors.library.caltech.edu/records/yh17a-hdm31Vorticity Generation by Contoured Wall Injectors
https://resolver.caltech.edu/CaltechAUTHORS:20101122-094355204
Authors: Waitz, Ian A.; Marble, Frank E.; Zukoski, Edward E.
Year: 1992
A class of contoured wall fuel injectors was designed to
enable shock-enhancement of hypervelocity mixing for
supersonic combustion ramjet applications. Previous
studies of these geometries left unresolved questions
concerning the relative importance of various axial vorticity
sources in mixing the injectant with the freestream. The
present study is a numerical simulation of two generic fuel
injectors which is aimed at elucidating the relative roles of
axial vorticity sources including: baroclinic torque through
shock-impingement, cross-stream shear, turning of
boundary layer vorticity, shock curvature, and diffusive flux.
Both the magnitude of the circulation, and the location of
vorticity with respect to the mixing interface were
considered. Baroclinic torque and cross-stream shear were
found to be most important in convectively mixing the
injectant with the freestream, with the former providing for
deposition of vorticity directly on the fue1/air interface.https://authors.library.caltech.edu/records/pg1n6-a6682A Systematic Experimental and Computational Investigation of a Class of Contoured Wall Fuel Injectors
https://resolver.caltech.edu/CaltechAUTHORS:20101130-132815278
Authors: Waitz, Ian A.; Marble, Frank E.; Zukoski, Edward E.
Year: 1992
The performance of a particular class of fuel injectors for
scramjet engine applications is addressed. The contoured
wall injectors were aimed at augmenting mixing through
axial vorticity production arising from interaction of the
fueVair interface with an oblique shock. Helium was used to
simulate hydrogen fuel and was injected at Mach 1.7 into a
Mach 6 airstream. The effects of incoming boundary layer
height. injector spacing, and injectant to freestream pressure and velocity ratios were investigated. Results from threedimensional flow field surveys and Navier-Stokes
simulations are presented. Performance was judged in
terms of mixing, loss generation and jet penetration.
Injector performance was strongly dependent on the
displacement effect of the hypersonic boundary layer which
acted to modify the effective wall geometry. The impact of
the boundary layer varied with injector array spacing.
Widely-spaced arrays were more resilient to the detrimental
effects of large boundary layers. Strong dependence on
injectant to free stream pressure ratio was also displayed.
Pressure ratios near unity were most conducive to losseffective mixing and strong jet penetration. Effects due to variation in mean shear associated with non-unity velocity ratios were found to be secondary within the small range of values tested.https://authors.library.caltech.edu/records/1s1ng-kxc10Rayleigh Scattering Measurements of Shock Enhanced Mixing
https://resolver.caltech.edu/CaltechAUTHORS:20101129-084947901
Authors: Budzinski, John M.; Zukoski, Edward E.; Marble, Frank E.
Year: 1992
This investigation was concerned with the nuxmg
which occurs after the unsteady interaction of a shock
wave with a laminar jet of helium. The jet of helium was
injected normal to the direction of the propagation of the
shock. The primary diagnostic, planar Rayleigh
scattering, had sufficient spatial and temporal resolution
to resolve the smallest diffusion scales present and
allowed helium mole fractions to be measured in twodimensional
planes normal to the original jet flow
direction. The amount of molecular mixing was
evaluated with a mass distribution function at increasing
times after the shock interaction. The total masses of
helium contained in regions where the molar
concentration of helium was at least 30% and 50% were
also calculated. The shock Mach number was varied, and
the effect of a reflected shock was studied. It was found
that shock interactions can significantly increase the
mixing between the air and helium. A rough collapse of
the mixing data occurs when time is normalized by the jet
radius divided by the change in velocity of the air behind
the shock. An increase in the enhancement of mixing
occurred after the interaction with the reflected shock.https://authors.library.caltech.edu/records/j63pr-mrn65An experimental study of the flow after shock interactions with cylindrical helium inhomegeneities
https://resolver.caltech.edu/CaltechAUTHORS:20110204-100414048
Authors: Budzinski, J. M.; Marble, F. E.; Zukoski, E. E.
Year: 1995
A shock traveling in air interacts with a laminar jet of
helium flowing normal to the direction of shock
propagation. Planar laser Rayleigh scattering is used
to study the deformation and motion of the originally
circular jet cross-section. The velocity of the jet
before the shock interaction is much less than the
velocities generated by the shock wave. Thus, the
helium jet serves to create a cylindrical bubble of a
lighter density gas imbedded in a heavier one. Four
different shock Mach numbers (1.066, 1.14, 1.5, and
2.0) are studied. Two different jet/air density ratios
are examined by using pure helium in the jet in one
case, and a mixture of airlhelium in the other. After
the shock interaction, a vortex pair forms from the baroclinically generated vorticity. The experiments
measure the velocity of the helium relative to the
surrounding air, the spacing between the vortex
cores, and the circulation of the vortices.
Experiments viewing the reflected shock interaction
are also performed. Excellent agreement is found
with previous computational studies.https://authors.library.caltech.edu/records/6gmq7-q4c82Non-Uniform Flow in Multistage Axial Compressors
https://resolver.caltech.edu/CaltechAUTHORS:20151111-165514089
Authors: Marble, Frank E.
Year: 2015
It has been suggested by the author that some aspects of
severely distorted flow into multistage compressors may be examined utilizing an integral technique. The general idea of the proposed technique is clear enough; the appropriate equations of motion and energy are integrated peripherally and radially, using reasonable assumptions for the distributions of velocity and thermodynamic
properties, and thereby reduced to ordinary non-linear differential equations for the parameters that describe the distributions. The questions that arise are whether the cascade characteristics may be described appropriately over wide variations of inlet angle, including stall, and whether the profiles may be characterized by a
sufficiently small number of parameters to make the technique attractive.
The present paper examines a specific example of distorted
inlet flow through the two-dimensional annulus of a multistage compressor which can be solved completely. It is shown that the essential features of this exact solution, including stall, may be described by a two-parameter family of profiles and that an integral technique,
utilizing these elementary profiles, will yield essentially the same results. While it is not clear that comparable success would hold for the three-dimensional problem, the results confirm the contention that the two-dimensional problem may be treated with acceptable
accuracy by an integral technique.https://authors.library.caltech.edu/records/hbw48-1ky59