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A Caltech Library Repository Feedhttp://www.rssboard.org/rss-specificationpython-feedgenenThu, 30 Nov 2023 19:45:35 +0000A Study of Methods to Increase the Lift of Supersonic Airfoils at Low Speeds
https://resolver.caltech.edu/CaltechETD:etd-12082008-083805
Authors: Pollock, Albert David; Reck, Floyd Francis
Year: 1947
DOI: 10.7907/J3NG-DD26
This thesis presents a study of the problem of improving the lift characteristics of a supersonic wing at low speeds. Trailing edge split flaps, nose flaps, and boundary layer control were investigated singularly and together using the optimum configuration of each.
Results indicate that the nose flap has an appreciable effect on preventing separation and thus increasing the lift. Split flaps give an increment of lift as would be expected. The boundary layer control consisted of blowing a sheet of high velocity air back over the top surface of the wing with very definite improvements of the lift and drag characteristics.
The work on the blowing technique, it is suggested, indicates sufficient promise to warrant much further study. The relatively large increment of lift that can be attributed to the prevention of flow separation at high angles of attack suggests that such boundary layer control could be used to improve controlability and to delay the stall, particularly tip stall, of high speed aircraft with very large sweep back angles.
https://thesis.library.caltech.edu/id/eprint/4877An Investigation of Detached Shock Waves
https://resolver.caltech.edu/CaltechTHESIS:09152014-090058146
Authors: Marschner, Bernard Walter
Year: 1948
<p>This investigation demonstrates an application of a
flexible wall nozzle for testing in a supersonic wind tunnel.
It is conservative to say that the versatility of this nozzle
is such that it warrants the expenditure of time to carefully
engineer a nozzle and incorporate it in the wind tunnel as a
permanent part of the system. The gradients in the test section
were kept within one percent of the calibrated Mach number, however,
the gradients occurring over the bodies tested were only
± 0.2 percent in Mach number.</p>
<p>The conditions existing on a finite cone with a vertex
angle of 75° were investigated by considering the pressure distribution
on the cone and the shape of the shock wave. The
pressure distribution on the surface of the 75° cone when based
on upstream conditions does not show any discontinuities at the
theoretical attachment Mach number.</p>
<p>Both the angle of the shock wave and the pressure distribution
of the 75° cone are in very close agreement with the theoretical
values given in the Kopal report, (Ref. 3).</p>
<p>The location of the intersection of the sonic line with
the surface of the cone and with the shock wave are given for the
cone. The blocking characteristics of the GALCIT supersonic
wind tunnel were investigated with a series of 60° cones.</p>https://thesis.library.caltech.edu/id/eprint/8657An Experimental Investigation of Detached Shock Waves
https://resolver.caltech.edu/CaltechETD:etd-12232008-153257
Authors: Altseimer, John H.
Year: 1948
DOI: 10.7907/W6SF-EG14
<p>Using the Galcit 2 1/2 inch Supersonic Tunnel equipment, surface pressure measurements were taken on cylindrical models with cone-shaped noses. The total included angle of the cones was 75°. The Mach number range covered was from 1.41 to 1.99 which for 75° cones covers most of the range from detached to attached shock waves. Stagnation pressure measurements were also taken by means of blunt cylindrical models.</p>
<p>Schlieren pictures were taken of the 75° cones and blunt models throughout the velocity range. Additional pictures of shock waves 45°, 60°, and 90° cones at some of the velocities used were also obtained. From the Schlieren pictures the traces of the shock waves were obtained and plotted on graph paper. Shock wave angles vs. distance from the model centerlines are also presented as well as the distribution of "M[subscript 2]", the Mach number directly behind the wave.</p>
<p>Appendix #1 shows "Blocking" data for the 2 1/2 inch GALCIT Tunnel.</p>
<p>Hand adjusted flexible nozzle walls were used for most of the work. Satisfactory flow through the tunnel was obtained with these nozzle blocks. However, design improvements are possible and Appendix #2 recapitulates the problems encountered and offers suggestions for future use.</p>https://thesis.library.caltech.edu/id/eprint/5132Wind Tunnel Investigation of a Supersonic Tailless Airplane at Low Subsonic Speed
https://resolver.caltech.edu/CaltechETD:etd-12192008-103750
Authors: Jensen, Arnold Axtell; Koerner, Warren Gottlieb
Year: 1948
DOI: 10.7907/KFRK-AF65
An investigation was made in the Caltech-Merrill low speed wind tunnel at Pasadena City College to determine the lift and moment characteristics of a sweptback wing and a comparable delta wing, both with a 65° sweptback leading edge and a double wedge symmetrical airfoil section. Both wings were tested with and without a fuselage. Leading edge flaps and slats, trailing edge plain flaps, split flaps, Fowler type flaps, and fillets were tried to determine their effects on these characteristics. The complete airplane was designed with the idea that it should be a tailless airplane.
The results showed unfavorable longitudinal static stability characteristics which could be improved but which could never be completely overcome at the stall when the wings were tested with the fuselage. A horizontal tail surface was necessary for longitudinal static stability at the stall but proved ineffective at the lower angles of attack.
The maximum lift coefficients for both wings of about 1.3 were higher than for a two dimensional double wedge airfoil section of approximately 0.8. The angles of attack at which these were reached were about twice as high as for the two dimensional section.
Tuft surveys showed the formation of two strong vortices from the leading edge of both wings first appearing at an angle of attack of approximately 10°. These vortices separated from the upper surface of the wing before reaching the trailing edge.
Comparison of results for the two wings indicated that the discontinuity of the trailing edge at the root of the sweptback wing was detrimental to the maximum lift.
There was an optimum deflection of the trailing edge split flap as a high lift device.
On the delta wing alone the plain flaps were very effective in increasing the maximum lift while the split flaps were ineffective.
Ailerons on the sweptback wing wore effective at all angles of attack through the stall.https://thesis.library.caltech.edu/id/eprint/5076Study of Vortex Shedding as Related to Torsional Oscillations of a Thin Airfoil
https://resolver.caltech.edu/CaltechETD:etd-12162008-142937
Authors: Magnus, Richard Jeffrey
Year: 1948
DOI: 10.7907/NAKT-DG24
This report covers an experimental investigation of the relationship between the vortex shedding frequency and self excited torsional oscillation frequency for a thin airfoil. The work consisted of measurements of velocity fluctuations in the airstream in the vicinity of a wing model mounted in a wind tunnel so that it could oscillate about the wing axis. The velocity fluctuation measurements were made with the Wing restrained and with the wing oscillating at various angles of attack and wind velocities.
Two distinct types of oscillations were found. One type was self sustaining and increased in amplitude with increasing wind velocity while the other type stopped for velocities beyond some critical value.https://thesis.library.caltech.edu/id/eprint/5027An Investigation of Vortex Shedding as Related to the Self-Excited Torsional Oscillation of an Airfoil
https://resolver.caltech.edu/CaltechETD:etd-02102009-131702
Authors: Chuan, Raymond Lu-Po
Year: 1948
DOI: 10.7907/6JQH-FS81
<p>This report covers the results of the experimental investigation of the self-excited torsional oscillation of a NACA 0006 airfoil suspended elastically. The relationship between the torsional oscillation and the shedding of vortices was investigated for this airfoil.</p>
<p>Two types of oscillation phenomena were found in the investigation. One type, exhibited by cases with angles of attack just above stall, persisted with increasing velocity without reaching any apparent limit within the range of velocity attainable in the present wind-tunnel. The other type, exhibited by cases with higher angles of attack, only showed self-excited oscillations in a certain range of velocity, the range decreasing with increasing angle of attack.</p>https://thesis.library.caltech.edu/id/eprint/600Sounding Rocket Performances Analysis
https://resolver.caltech.edu/CaltechETD:etd-01232009-152718
Authors: Burke, James Donahue
Year: 1949
DOI: 10.7907/XW5Q-1Y06
An investigation of the effects of basic design parameters on the performance of a single-stage sounding rocket was made. The tail area required for static stability at various supersonic flight Mach numbers was determined, and the drag coefficient for various configurations was calculated. The general equations of motion for flight in vacuum were integrated for several cases to show the effect of varying specific impulse of fuel, propellant weight ratio, and burning time. An approximate solution of the equations for trajectories in air was obtained by the method of stepwise numerical integration.https://thesis.library.caltech.edu/id/eprint/314A Wind Tunnel Investigation of the Aerodynamic Characteristics of High-Lift Devices on Supersonic Wings at Low Subsonic Speed
https://resolver.caltech.edu/CaltechETD:etd-01272009-151720
Authors: Blenkush, Philip George
Year: 1949
DOI: 10.7907/N6XA-GR06
An investigation was made to determine the effects of various full- and partial-span high-lift devices on the lift, drag, and pitching moment characteristics of a straight wing and a highly sweptforward wing, both having an aspect ratio of 1.72 and a thin double wedge symmetrical airfoil section. Split, extended leading edge, and extended trailing edge flaps were tested at various deflection angles on each wing alone and also on wing-fuselage combinations. In addition, plain leading edge flaps extending over a portion of the outboard span were tested on the sweptforward wing. Tuft surveys were made on typical model configurations to determine the change in stall pattern due to variation in the angle of attack.
From the results of the tests it was found that a given flap configuration produced approximately equal increments of maximum lift on the sweptforward wing alone and on the same wing combined with a fuselage. Comparison of the results obtained with the straight wing showed that the maximum lift increments were lower for a given flap used on wing plus fuselage than those obtained with the same flap on the basic straight wing. When considered from the maximum lift standpoint, the extended trailing edge flap was superior to either split or extended leading edge flap. Reduced span flaps were not as effective on the sweptforward wing, as the comparable full span flap. In the case of the straight wing alone, reduced span split and extended trailing edge flaps produced higher maximum lifts than did the corresponding full span configuration. This effect was also observed with the split flap tests on the straight wing combined with the fuselage.
The experimental work was performed in the Cal Tech-Merrill wind tunnel located on the Pasadena City College campus.https://thesis.library.caltech.edu/id/eprint/383Low Speed Wind Tunnel Investigation of High Lift Devices on a 65° Swept-Back Supersonic Wing of 3.44 Aspect Ratio
https://resolver.caltech.edu/CaltechETD:etd-02242009-092302
Authors: Thomas, John William
Year: 1949
DOI: 10.7907/TYRW-TZ93
NOTE: Text or symbols not renderable in plain ASCII are indicated by [...]. Abstract is included in .pdf document.
A low speed investigation was made in the Cal Tech-Merrill Wind Tunnel at Pasadena City College of a 3.44 aspect ratio wing, having a 65° swept-back leading edge and a symretrical double-wedge airfoil section, to determine the characteristics and effectiveness of various high-lift devices. The investigation included tests of trailing edge split flaps and extended split flaps, leading edge flaps and slats, and combined configurations. The tests were carried out using the wing with and without the fuselage and tail surfaces.
The results showed that the split flaps increased the lift over the lower regions of the angle of attack only. The extended split flaps were more effective and increased the lift over the whole range. The nose flaps tested did not increase the lift when used by themselves, but the combination of a nose flap with the fuselage and extended split flap produced the greatest lift. The addition of the fuselage added to the lift by increasing the slope of the lift curve. The addition of the tail surface gave additional increments in lift in the upper regions of angle of attack. The maximum lift coefficient, [...], normally occurred at an angle of attack of 38°.
The extended split flaps produced large variations in pitching moment, but had stabilizing tendencies except where there were irregularities in the lift curve.https://thesis.library.caltech.edu/id/eprint/745Flow Field Around a Finite Cone with Shock
https://resolver.caltech.edu/CaltechETD:etd-02182009-110248
Authors: MacKinnon, Neil Allan
Year: 1949
DOI: 10.7907/XR4M-9R51
An experimental investigation was made to determine the characteristics of the flow over the surface of a 70° cone and at the shock wave for values near the detachment Mach number. The purpose of this investigation was to compare the experimental results obtained with theoretical values.
Tests were made in the GALCIT 2.5" Supersonic Wind Tunnel on a 70° cone at zero angle of attack for five different free stream Mach numbers: 1.49, 1.630, 1.694, 1.86, 1.997.
It was found that theory gives close agreement with experimental results.
This investigation was conducted jointly with Mr. Vincent Muirhead at the California Institute of Technology, Pasadena, California.https://thesis.library.caltech.edu/id/eprint/675Wind Tunnel Investigation of the Aerodynamic Characteristics of High Lift Devices on Supersonic Wings at Low Subsonic Speed
https://resolver.caltech.edu/CaltechETD:etd-01232009-104437
Authors: Densmore, James Edward
Year: 1949
DOI: 10.7907/Z0Y9-J658
An investigation was made to determine the effect of various high lift devices on a highly sweptforward wing with a leading edge sweep angle of 55 degrees and on a straight wing, both wings having an aspect ratio of 1.72 and the same span.
Experimental tests were made in the Cal Tech - Merrill low speed wind tunnel at Pasadena City College on both types of wing with and without fuselage. High lift devices investigated were extended leading edge flaps, plain leading edge flaps, extended trailing edge flaps, and split flaps. Both 70 per cent and full span configurations were used in each case except for the plain leading edge flaps. All force data was reduced to standard non-dimensional lift, drag, and pitching moment coefficients, and the results presented in standard graphical form. In addition, photographs of tuft surveys were made for typical configurations.
The maximum lift coefficient obtained from the basic wings was approximately the same for both, but the angle of attack for maximum lift was appreciably lower for the straight wing than for the sweptforward wing.
The addition of the fuselage increased the maximum lift coefficient of both the basic wings and of the sweptforward wing with flap configurations, but the fuselage was detrimental to the lift for the straight wing with flap configurations.
Straight wing configurations gave respectively higher maximum lift coefficients and larger lift curve slopes than on the comparable sweptforward wing models. All high lift devices investigated were more effective on the straight wing than on the sweptforward wing.
Extended trailing edge flaps were the most effective of the flaps investigated in increasing the maximum lift, but gave the largest negative increase in the pitching moments.
On the straight wing the 70 per cent span split and 70 per cent span extended trailing edge flaps at optimum flap deflection angles gave a higher maximum lift coefficient than the full span flaps at the optimum flap deflection. When the straight wing was mounted on the fuselage, this effect was true only for the split flaps.https://thesis.library.caltech.edu/id/eprint/311Winged Rocket Performance Analysis
https://resolver.caltech.edu/CaltechETD:etd-02052009-092504
Authors: Shonerd, David Edwin
Year: 1949
DOI: 10.7907/34ZH-KG02
A method is presented for estimating the effects of various parameters on the performance of a winged rocket. A program for studying three specific parameters, i.e., wing areas, reduced thrust cruising programs, and trajectory climb angles, and their effect on the horizontal range of a winged rocket is presented. Complete calculations are carried out for one combination of these parameters.
An analysis of the Lift, Drag and Stability characteristics of a long, slender rocket with trapezoidal fins and wings is made. The stabilizing effectiveness of delta and trapezoidal fins is compared.
Simplified approximate methods of integrating the trajectory equations by step-by-step method are presented.https://thesis.library.caltech.edu/id/eprint/511Flow Field Around a Finite Cone with Shock
https://resolver.caltech.edu/CaltechETD:etd-01282009-092339
Authors: Muirhead, Vincent Uriel
Year: 1949
DOI: 10.7907/7GMA-0V07
The objective of this investigation was to study the flow field in the immediate vicinity of a finite cone and to compare the results with analytical values for an infinite cone. Pressure distribution over the surface of a 70 degree cone and the general characteristics of the shock wave were investigated. The tests were conducted at five Mach numbers covering the four regimes of flow. Particular attention was given to the conditions at the apex of the cone. The locations of the intersection of the sonic line with the surface of the cone and with the shock wave were determined.
In general the tests demonstrate that at the apex of a finite cone the pressure and the shock wave angle closely approach the values predicted by analytical methods in the four regimes of flow.
The tests were conducted in the GALCIT 2.5" Supersonic Wind Tunnel.https://thesis.library.caltech.edu/id/eprint/392An Investigation of Blowing as a Method of Increasing the Maximum Lift of a Double-Wedge Airfoil
https://resolver.caltech.edu/CaltechETD:etd-01272009-155604
Authors: Bacon, John William
Year: 1949
DOI: 10.7907/NT63-HE54
This report presents the results of an investigation of boundary layer energization by blowing as a method of increasing the maximum lift coefficient for supersonic airfoils operating at low subsonic speeds.
The model used was a double wedged airfoil with end plates and a spanwise slot located at the 15%, 25% or 35% chord position. Air was blown back over the wing from this slot producing a considerable increase in maximum lift for most configurations. Blowing proved to be most effective in increasing the maximum lift when used with the nose flap, fairly effective with combined flaps and wing alone, but ineffective with the split flap.https://thesis.library.caltech.edu/id/eprint/385Low Speed Wind Tunnel Investigation of High Lift Devices on a 65° Swept-Back Supersonic Wing of 3.44 Aspect Ratio
https://resolver.caltech.edu/CaltechETD:etd-02122009-093718
Authors: Marx, Michael Ferdinand
Year: 1949
DOI: 10.7907/2B48-J790
A low speed survey was conducted on a 3.44 aspect ratio wing having a 65° swept-back leading edge and double wedge symmetrical airfoil section to obtain information as to the effectiveness of various high-lift devices. These devices included trailing edge split and extended split flaps, leading edge split flaps, slats and combined configurations. Tests were carried out on the wing with and without the fuselage and horizontal tail surface.
The split flaps increased the lift over the lower ranges of angle of attack only. The extended split flaps increased the lift over the whole angle of attack range. Hose flaps showed practically no gain over any of the range when used by themselves. However, when combined with the trailing edge split flaps in the wing-fuselage configuration, the optimum maximum lift conditions were obtained. Addition of the fuselage and horizontal tail surfaces each produced considerable increments of lift.
In all configurations except leading edge flaps in the inboard position undesirably large negative pitching moments resulted. However, they had stabilizing tendencies except where there were irregularities in the lift curves.https://thesis.library.caltech.edu/id/eprint/620A Study of Detached Shock Waves in Two-Dimensions
https://resolver.caltech.edu/CaltechETD:etd-02022009-081306
Authors: Alperin, Morton
Year: 1950
DOI: 10.7907/NW30-HY03
<p>The present report contains results of an experimental and theoretical investigation of the detached shock wave phenomenon. The experimental phase of this study was actually carried out at the Jet Propulsion Laboratory at California Institute of Technology, in a two-dimensional wind tunnel which is briefly described in Section I.</p>
<p>Section II contains a description of the experiments on circular cylinders. The circular cylinder was used in this series of tests primarily because of its simplicity. The investigation discussed in II-1 required a large variation of model shapes and would have required much more time had it been based on a more complicated body shape. In addition to data on the shock wave position and shape, the pressure distribution was also obtained at M=1.546 for a two-dimensional circular cylinder. From this pressure distribution, the drag was calculated.</p>
<p>Although the theoretical knowledge of flow involving detached shock waves is in a rather primitive state, a review of the existing theoretical work and comparison with experimental data is made in section III.</p>
<p>In section IV a method is presented for finding the stream function or velocity potential for the subsonic region behind the detached shock wave. This method depends upon the hypothesis that the flow can be considered to be irrotational in this region without introducing a serious error. The results appear to be in good agreement with the experiments although the example carried out does not apply strictly to the circular cylinder body shape used in the experiments.</p>
<p>A general discussion of the existing theories and their comparison with experimental data is presented in section V.</p>https://thesis.library.caltech.edu/id/eprint/471Acoustical Airspeed Indicators
https://resolver.caltech.edu/CaltechETD:etd-02022009-083911
Authors: Barish, David Theodore
Year: 1950
DOI: 10.7907/FR7P-K372
Some of the problems associated with the applications of acoustical devices for the determination of airstream characteristics are considered in this study. The velocity and pressure fields for both point sound sources and finite sound sources in both subsonic and supersonic flow are discussed, with a view toward using sound waves for the determination of velocity, Mach number , temperature, and other properties of a flow.
The experimental investigations included the measurement of the spectra of ultra-audio-pressure pulsations, both static and total, in the small C.I.T. supersonic tunnel and also in the C.I.T. hypersonic tunnel. A broad range of Mach numbers and a variety of operating conditions were covered. The development of a modified Hartmann sound generator is described, and measurements of the sound field from this device in supersonic flow are included.https://thesis.library.caltech.edu/id/eprint/472Experimental Investigation of Detached Shock on a 70° Cone at Various Angles of Attack
https://resolver.caltech.edu/CaltechETD:etd-02042009-100038
Authors: Bryan, William Cleveland
Year: 1950
DOI: 10.7907/XAZW-8051
The results of an experimental investigation into the variation of the shock wave shape and the extent of the subsonic region behind the shock with varying angle of attack at various Mach numbers are summarized. All tests are made on a finite 70° cone. The main interest is centered on those Mach numbers which are low enough to produce detached shocks or for which the possibility of detachment exists for increasing angles of attack.
The angles of attack for which the investigation is made are 0°, 3°, 6°, 9°, 12° and 15°. The Mach numbers considered are 1.438, 1.544, 1.584, 1.857, 1.986, and 3.01.
It is found that, for increasing angles of attack at constant Mach numbers, the extent of the subsonic region increases behind the lower shock and decreases behind the upper shock. A subsonic region appears at increased angle of attack for two of the Mach numbers for which the shocks are initially attached. There is also quite definite interaction between the upper and lower portions of the shock which tends to inhibit both the appearance of a subsonic region after the shock on the one hand and the disappearance of it on the other.https://thesis.library.caltech.edu/id/eprint/498Flexible Plate Supersonic Wind Tunnel Flow Correction to Account for Plate Elastic Properties
https://resolver.caltech.edu/CaltechETD:etd-02242009-093505
Authors: Chase, Patrick Stanley
Year: 1950
DOI: 10.7907/EC0T-P917
A method of correcting the static pressure distribution in the working section of a supersonic wind tunnel with flexible nozzle walls is investigated, in order to account for the discrepancy between the theoretical nozzle wall shape and the actual wall shape due to its elastic properties.
The method of correction is described and an example for an existing wind tunnel is carried out. It is shown that the method of correction is not successful due to the inadequate accuracy of the data from physical measurements and due to the nature of the equations which must be solved. While the former difficulty could possibly be avoided, the latter is considered to be inherent in the method.
Suggestions and criticisms are offered and some results of the investigation regarding correction of the actual wind tunnel static pressure distributions are mentioned.https://thesis.library.caltech.edu/id/eprint/746Aerodynamic Characteristics of a Wedge and Cone at Hypersonic Mach Numbers
https://resolver.caltech.edu/CaltechThesis:09292023-233404349
Authors: DeLauer, Richard Daniel
Year: 1950
DOI: 10.7907/rndh-4a17
<p>The problem of predicting the aerodynamic characteristics of configurations at hypersonic Mach numbers has been unreliable due to the lack of experimental data.</p>
<p>By predicting the aerodynamic characteristics of a wedge and cone at Mach numbers from 2 to 12 by four different supersonic theories, a basis for future experimental comparison was provided.</p>
<p>An attempt was made to correlate the theoretical result of a 20° wedge and cone with wind tunnel test results of the same configuration. However, due to scheduling difficulties the experimental phase was not completed in time enough to be included in this report.</p>
<p>The theoretical results indicate that the hypersonic similarity solution gives close agreement with the exact solution for large Mach numbers. The linearized and second order theory deviates from the exact solution for Mach numbers greater than 3.</p>https://thesis.library.caltech.edu/id/eprint/16199Experimental Investigation of Detached Shock on a 70° Cone at Various Angles of Attack
https://resolver.caltech.edu/CaltechETD:etd-02122009-125456
Authors: Frick, Leo Francis
Year: 1950
DOI: 10.7907/D4TV-0W57
An experimental investigation was made to determine the variation of shock shape and extent of subsonic region behind the shock wave with angle of attack for a 70° cone at various Mach numbers. The main interest was centered on those Mach numbers which produced detached shock waves or for which the possibility of detachment at angle of attack existed.
The tests were conducted in the GALCIT 2.5" Supersonic Wind Tunnel at angles of attack of 0°, 3°, 6°, 9°, 12°, and 15° and Mach numbers of 1.438, 1.544, 1.584, 1.857, 1.986, and 3.01.
It was found that, with increasing angle of attack and constant Mach number, the subsonic region behind the shock wave increased in the lower portion and decreased in the upper portion. With increasing Mach number the subsonic region decreased for all angles of attack. Interaction between the upper and lower portions of the shock wave affects the subsonic region behind the wave, suppressing its appearance on the upper surface, retarding its disappearance from the lower surface.https://thesis.library.caltech.edu/id/eprint/621Aerodynamic Characteristics of a Wedge and Cone at Hypersonic Mach Numbers
https://resolver.caltech.edu/CaltechETD:etd-02112009-150520
Authors: Scherer, Lee Richard
Year: 1950
DOI: 10.7907/29PE-XX71
Up to the present time, the reliability of the determination of aerodynamic characteristics at hypersonic Mach numbers by theoretical calculations has been unknown due to the lack of experimental data. This report is the calculations of these characteristics by four different theories of a wedge and a cone over a range of Mach numbers from 2 to 12.
Correlation of these results with wind tunnel tests was not possible due to scheduling difficulties of the hypersonic wind tunnel; therefore, this report is designed to serve as the basis for comparison of future hypersonic experiments.
From correlation of the various theories it is found that the closest agreement to the exact theory at hypersonic speeds is given by the hypersonic similarity theory. Above Mach numbers of about 3, the first and second order theories deviate considerably from the exact theory.https://thesis.library.caltech.edu/id/eprint/608Effects of Impurities on the Supersaturation of Nitrogen in a Hypersonic Wind Tunnel
https://resolver.caltech.edu/CaltechETD:etd-03232009-152948
Authors: Arthur, Paul David
Year: 1952
DOI: 10.7907/PBSD-2V70
An experimental investigation was conducted to determine the effects of additives on the supersaturation of commercial bottled nitrogen expanded in a hypersonic nozzle. In particular, enough oxygen was added to duplicate air proportions. A stainless steel two-dimensional source-flow nozzle of one-inch width was used to conduct the tests.
Commercially pure nitrogen, expanded from room temperature and 8-1/3 atm. pressure, was found to supersaturate by approximately 18° K or 1.2 Mach number. The supersaturation of the nitrogen was decreased by the addition of impurities, and only a fraction of a percent of carbon dioxide or water vapor was required to eliminate completely all supersaturation. Addition of argon and oxygen was found to be much less effective in decreasing the supersaturation. For the synthetic air, the supersaturation was 16° K or 0.9 Mach number based on air vapor pressure values.
During the collapse of the supersaturated state, the static pressure gradually increased above the isentropic value because of the heat release of the condensing gas. As has been shown before, there was no evidence of condensation shock with nitrogen. The impact pressure was only slightly changed from the isentropic value by the presence of condensation in the flow. After the collapse of the supersaturated state, the flow approximated that of a condensation shock.
From these tests it is concluded that condensation of nitrogen, containing slightly more impurities than present in the commercial nitrogen, and of air of the same purity principally caused by foreign impurities, not by spontaneous self-nucleation.https://thesis.library.caltech.edu/id/eprint/1081The Effects of Air Condensation on Properties of Flow and their Measurement in Hypersonic Wind Tunnels
https://resolver.caltech.edu/CaltechETD:etd-03302009-061157
Authors: Grey, Jerry
Year: 1952
DOI: 10.7907/HFHE-HR77
Some of the fundamental problems encountered in the measurement of flow properties in condensing air have been discussed, and were investigated experimentally in the GALCIT 9 Hypersonic Wind Tunnel. The saturated expansion theory of flow in a condensing fluid as developed by Buhler was corroborated, and some of the unknown properties of the theory have been clarified by analysis of the wind tunnel tests. Several experimental techniques for the measurement of two-phase fluid properties were developed and the results were used to supply additional information concerning the nature of phenomena such as supersaturation and normal and oblique shock waves. No definite conclusions could be reached with regard to future application of condensed air wind tunnel data on the basis of these tests, but the subject of similarity parameters comparable to the Mach number in flow of a perfect gas has been discussed at some length.https://thesis.library.caltech.edu/id/eprint/1204Condensation of Air Components in Hypersonic Wind Tunnels: Theoretical Calculations and Comparison with Experiment
https://resolver.caltech.edu/CaltechETD:etd-01162007-152444
Authors: Buhler, Rolf Dietrich
Year: 1952
DOI: 10.7907/0J8Q-9Z82
The effect of condensation on the flow in hypersonic wind tunnels is bracketed by equilibrium saturated expansion and by instantaneous condensation. By calculation of shock waves with evaporation, direct comparison of theoretical and measured pressures is made possible. Satisfactory agreement between saturated expansion theory and experiment is obtained after the collapse of the supersaturated state.
The droplet growth theory (for free molecule regime) is reexamined, and a good approximate solution is obtained for the nonsteady case (i.e., rapidly changing vapor properties). Limits of validity of the quasi-steady theory are defined, and an upper limiting (zero growth) drop size given for expanding flow.
A simplified method is presented for calculating the pressure time history of the collapse of the supersaturated state in nozzles. From this, most effective nucleus sizes for given total mass of impurities are calculated. Thus the earliest possible collapse in a nozzle due to impurities is estimated theoretically for low impurity concentrations. The agreement of the predicted trend with experimental results in nitrogen appears to justify the assumed mechanism of the collapse, which is condensation on existing foreign nuclei formed upstream of the collapse.
https://thesis.library.caltech.edu/id/eprint/197Investigation of an Instability Phenomena Occurring in Supersonic Diffusors
https://resolver.caltech.edu/CaltechETD:etd-05122003-115303
Authors: Stoolman, Leo
Year: 1953
DOI: 10.7907/EG52-8V58
Experimental investigations of supersonic normal shock type diffusors have shown the existence of self-excited oscillations that occur as the internal mass flow is reduced somewhat below its maximum value. There is a lower bound of free stream Mach number (of the order of 1.8) below which no instability could be observed. However, as free stream Mach number was increased above this lower bound, instability occurred at increasing values of the internal mass flow. Also, the frequency at instability was of the order of the natural frequency of the internal duct acting as organ pipe.
First-order theoretical investigations of the above phenomena indicate that the instability may (in part) be interpreted as intrinsic, that is, independent of viscous effects at the duct inlet or within the diffusor. The fundamental cause of the instability is shown to be due to the nature of the oscillatory inlet flow conditions that occur as a consequence of the external compression from the shock wave to the inlet, and the type of reflections suffered at the shock wave by upstream traveling pressure waves.https://thesis.library.caltech.edu/id/eprint/1737Boundary Layer Temperature Recovery Factor on a Cone at Nominal Mach Number Six
https://resolver.caltech.edu/CaltechETD:etd-05122003-095635
Authors: Mackay, Douglas Severance
Year: 1953
DOI: 10.7907/8DEC-WE68
NOTE: Text or symbols not renderable in plain ASCII are indicated by [...]. Abstract is included in .pdf document.
An investigation was conducted to determine the temperature recovery factors for laminar boundary layer on a cone at free stream Mach numbers from 5.6 to 5.9. The investigation was conducted in the GALCIT 5" x 5? continuous-flow, closed-circuit wind tunnel (Leg No. 1). Two twenty degree cone models about three inches in length were used. One model was composed of a ceramic core with a thin (0.010? to 0.015?) steel surface, and the second was a hollow copper shell of 0.005? thickness.
One-phase and two-phase (condensation) flow conditions were investigated. Temperature recovery factors were determined from the data obtained from the tests conducted with one-phase airflows. The ratios of the temperatures recovered on the cone surface to the respective stagnation temperatures were computed from the data obtained in the two-phase airflow investigations and were compared with these ratios for the one-phase airflows.
The local temperature recovery factors for the laminar boundary layer were determined to be 0.844 ? 0.008 for Reynolds numbers from 2.1 x 10[superscript 4] to 5.4 x 10[superscript 5] . For this range of Reynolds numbers the recovery factor was found to be independent of the Reynolds number. The independence of the recovery factor on the Mach number was substantiated (by comparison with results of previous investigations at lower Mach numbers) for Mach numbers up to 5.9. The ratios of the temperature recovered on the cone to the stagnation temperature were found to be the same for one and two-phase airflows.
The square root of the Prandtl number evaluated at the mean of the temperatures of the cone surface for the various flow conditions investigated was found to be less than one per cent lower than the mean of the experimental temperature recovery factors.
The results of this investigation are in agreement with those of previous investigations at lower Mach numbers and, within the limits of experimental accuracy, verify theoretical solutions.https://thesis.library.caltech.edu/id/eprint/1729Direct Measurement of Laminar Skin Friction at Hypersonic Speeds
https://resolver.caltech.edu/CaltechETD:etd-04292003-112020
Authors: Eimer, Manfred
Year: 1953
DOI: 10.7907/MK2E-1P76
NOTE: Text or symbols not renderable in plain ASCII are indicated by [...]. Abstract is included in .pdf document.
A direct measurement of flat plate laminar local skin friction was undertaken at M = 5.8 in the GALCIT Hypersonic Wind Tunnel, Leg No. 1. A new balance particularly suited to the requirements of hypersonic experimentation was designed.
By means of the fluorescent lacquer technique for indicating boundary layer transition, unexpectedly high natural transition Reynolds numbers were observed. The several methods of forcing transition which were used were unsuccessful in producing transition at a Reynolds number of two million.
Observation of the leading edge shock wave and boundary layer by means of a schlieren system indicated that at M = 5.8 the shock wave and boundary layer are separated by appreciable distances wherever the boundary layer equations hold.
Skin friction force measurements were made at five Reynolds numbers in condensation-free flow. The discrepancies between the observed low values of C[subscript f][...]Re and existing laminar boundary layer calculations are reconciled by means of a qualitative analysis.
A theory describing the properties of the viscous boundary layer for flows with condensation in the free stream is presented. It is predicted that for a specified nozzle geometry and stagnation temperature, skin friction remains unchanged in the presence of moderate amounts of condensation, while heat transfer for a given wall temperature is affected by the presence of condensation which produces major changes in the adiabatic wall temperature.https://thesis.library.caltech.edu/id/eprint/1539Experimental Heat Transfer at Hypersonic Mach Number
https://resolver.caltech.edu/CaltechETD:etd-04222003-111626
Authors: DeLauer, Richard Daniel
Year: 1953
DOI: 10.7907/59SJ-TR03
<p>An experimental investigation was conducted in Leg 1 of the GALCIT 5 x 5 inch Hypersonic Wind Tunnel to determine the heat transfer coefficients of the laminar boundary layer on a cooled flat plate at a nominal Mach number of 5.8. As a consequence of the investigation, flat plate recovery factors were determined and the effect of condensation on heat transfer was noted. In addition qualitative results as to the laminar boundary layer transition and separation are also presented.</p>
<p>The tests were conducted with a ratio of wall temperature to free stream temperature (T<sub>w</sub>/T<sub>δ</sub>) of approximately 6.2; but under stagnation temperature conditions ranging from 200°F to 285°F. The stagnation pressure range of 60 psia to 115.5 psia provided a maximum Reynolds number of 2.1 x 10<sup>6</sup>.</p>
<p>A flat plate temperature recovery factor of .858 ± .004 was determined, and it was concluded that the temperature recovery factor range of Mach number independence could be extended to a Mach number of 5.8. The independence of the recovery factor on Reynolds number up to the beginning of the laminar boundary layer transition was also substantiated.</p>
<p>The heat transfer coefficients were obtained for a negative temperature gradient over a considerable portion of the plate. The effect of these gradients produced values considerably higher than would be expected for an isothermal surface. These results, when related the constant temperature case by a theoretical calculation, were in good agreement, with the theoretical results and the results of a friction investigation carried out at the same Mach number. The accuracy of the results was estimated to be ±10% from a value of Nu/Re<sup>1/2</sup>Pr<sup>1/3</sup> - .285.</p>
<p>There was no apparent effect on the heat transfer coefficient by condensation, but the adiabatic wall temperature appeared to be 2% lower than for the condensation free flow. Due to a step increase in thickness of the model at the ten inch station, the shock wave-boundary layer interaction appears to produce laminar boundary layer transition at a Reynolds number of 1.3 x 10<sup>6</sup>, and upon reducing the Reynolds number further, the transition point is subjected to an adverse pressure gradient which results in a boundary layer separation.</p>https://thesis.library.caltech.edu/id/eprint/1447Impact Pressure and Total Temperature Interpretation at Hypersonic Mach Numbers
https://resolver.caltech.edu/CaltechETD:etd-12052003-095631
Authors: Quiel, Norwald Richard
Year: 1954
DOI: 10.7907/SMHA-4160
An experimental investigation was undertaken at a nominal Mach number of 5.6 in the GALCIT Hypersonic Wind Tunnel, Leg No. 1. The first phase was an investigation of the viscous effects on measured impact pressures. The second was an investigation of the temperature recovery characteristics of a singly shielded total-temperature probe.
Experimental results are presented for a straight, sharp-lipped, cylindrical, impact-pressure probe and for a flattened-end probe. Impact-pressure data were obtained for a Reynolds number range from 425 to 8,000, where the Reynolds number was based on free stream conditions and the impact probe outside diameter, The data show that the Rayleigh equation requires corrections for viscous effects at Reynolds numbers less than 6,000 for the circular sharp-tipped probe and less than 4,000 for the flattened-end probe. The viscous effects increase with decreasing Remolds numbers. At a Reynolds number of 425, the measured impact pressure is approximately 2.5 per cent lower than that predicted by the Rayleigh equation. It was concluded that the viscous effects were dependent on Mach number as well as Reynolds number.
Temperature recovery factors for the total-temperature probe were obtained throughout a Reynolds number range from 30,800 to 213,000, where the Reynolds number was based on the probe entrance outside diameter and the free stream conditions. An analysis of suitable parameters with which to present the data is included together with the experimental data. For a limited range of total temperatures, a single temperature recovery calibration curve was obtained when the Reynolds number was used as a parameter. The data show that the temperature recovery factor of the total temperature probe decreases with decreasing Reynolds numbers.https://thesis.library.caltech.edu/id/eprint/4789Base Pressure and Static Pressure for a Cone-Cylinder at a Nominal Mach Number of 5.8
https://resolver.caltech.edu/CaltechETD:etd-12152003-095830
Authors: Harkins, William Douglas
Year: 1954
DOI: 10.7907/P0K8-XG38
An experimental investigation was made in the GALCIT Hypersonic Wind Tunnel Leg No. l to determine the base pressure and static pressure on a cone-cylinder at a nominal Mach number of 5.8 in both one-phase and two-phase flow.
The scope of the investigation was a determination of interference data necessary for proper evaluation of base pressure results, investigation of the effect of Reynolds number on base pressure, and a comparison of experimental and theoretical static pressure distribution on a cone-cylinder.
As has been noted by other investigators, viscous effects in hypersonic flow were quite pronounced and demonstrated the increased non-linearity of the problems in hypersonic flow.https://thesis.library.caltech.edu/id/eprint/5001Impact Pressure and Total Temperature Interpretation at Hypersonic Mach Number
https://resolver.caltech.edu/CaltechETD:etd-12182003-100211
Authors: Graves, Jack Carl
Year: 1954
DOI: 10.7907/V04Q-TF29
Results are presented of an experimental investigation of impact-pressure and total-temperature interpretation at a nominal Mach number of 5.6. The data indicate that the Rayleigh equation, which assumes non-viscous flow, requires correction at low free-stream Reynolds numbers. These viscous effects are detected at Reynolds numbers (based on impact-probe diameter) as high as 6000, and they continue to increase with decreasing Reynolds numbers. At the low pressure limit of the facilities used in this investigation, the maximum viscous correction is 2.5 percent for a Reynolds number of 425.
The calibration curves for the recovery factor of a total-temperature probe are given, plus an analysis of suitable parameters with which to present this information. For the limited range of total temperatures of 200[degrees]F to 260[degrees]F, and a nominal Mach number of 5.6, single calibration curves are shown using either the free-stream Reynolds number, or the Nusselt number of the flow inside the probe (based on thermocouple wire diameter) as parameters.https://thesis.library.caltech.edu/id/eprint/5039Laminar Boundary Layers on Slender Bodies of Revolution in Axial Flow
https://resolver.caltech.edu/CaltechETD:etd-01142004-155216
Authors: Mark, Richard Muin
Year: 1954
DOI: 10.7907/4H6Z-K569
An exact similar solution of the modified boundary layer equations has been obtained for the axial incompressible flow past paraboloids of revolution. It has been shown that the usual boundary layer assumptions are justified and that the local skin friction increases as the boundary layer thickness becomes large compared with the body radius.
An approximate method for obtaining the local skin friction on arbitrary slender bodies of revolution in axial incompressible flow has been developed. A comparison of the approximate results with the exact solutions for paraboloids of revolution and circular cylinders shows good agreement.
The existence of energy integrals of the modified compressible boundary layer equations is established. Similarity of the governing equations for the axial compressible flow past paraboloids of revolution has been shown; for the same bodies, a hypersonic similarity law is deduced.
An approximate method for obtaining the local skin friction on arbitrary slender insulated bodies of revolution in axial compressible flow has been developed. The results show that compressibility counterbalances the rise in local skin friction due to curvature at high Reynolds numbers (based on a characteristic length of the body) and increases the local skin friction at sufficiently low Reynolds numbers.
Velocity profiles on a slender ogive-cylinder have been obtained experimentally at a Mach number of 5.8 and at different Reynolds numbers. The results indicate a curvature effect when compared with flat plate results.https://thesis.library.caltech.edu/id/eprint/168Transition Studies and Skin Friction Measurements on an Insulated Flat Plate at a Hypersonic Mach Number
https://resolver.caltech.edu/CaltechETD:etd-01072004-114554
Authors: Korkegi, Robert Hani
Year: 1954
DOI: 10.7907/X3B9-VW73
NOTE: Text or symbols not renderable in plain ASCII are indicated by [...]. Abstract is included in .pdf document.
An investigation of transition and skin friction on an insulated flat plate, 5 x 26 inches, was made in the GALCIT 5 x 5 inch Hypersonic Wind Tunnel, Leg No. 1, at a nominal Mach number of 5.8.
The phosphorescent lacquer technique was used for transition detection and was found to be in good agreement with total-head rake measurements along the plate surface and pitot boundary layer surveys. It was found that the boundary layer was laminar at Reynolds numbers of at least 5 x 10[superscript 6]. It was also observed that transverse contamination due to the turbulent boundary layer on the tunnel sidewall originated far downstream of the flat plate leading edge at Reynolds numbers of 1.5 to 2 x 10[superscript 6], and spread at a uniform angle of 5 1/2[degrees] compared with 9 1/2[degrees] in low speed flow.
The effect of two-dimensional and local disturbances was investigated. The technique of air injection into the boundary layer as a means of stimulating transition was extensively used. It was observed that, although the onset of transition occurred at Reynolds numbers down to 10[superscript 6], a fully developed turbulent boundary layer was not obtained at Reynolds numbers much below 2 x 10[superscript 6] regardless of the amount of air injected.
A qualitative discussion of these results is given with emphasis on the possibility of a greater stability of the laminar boundary layer in hypersonic flow than at lower speeds.
Direct skin friction measurements were made by means of the floating element technique incorporating a null system using chain loading, over a range of Reynolds numbers (based on distance from leading edge) from 10[superscript 6] to 4 x 10[superscript 6]. Without artificial tripping, the boundary layer was verified as being laminar over the complete range. With air injection, turbulent shear was obtained only for Reynolds numbers greater than 2 x 10[superscript 6] , this value being in good agreement with earlier results of this investigation. The turbulent skin friction coefficient was found to be approximately 0.40 of that for incompressible flow for a constant value of R[subscript theta], and 0.46 for an effective Reynolds number between 5 and 6 x 10[superscript 6].https://thesis.library.caltech.edu/id/eprint/50The Hypersonic Shock Tube
https://resolver.caltech.edu/CaltechETD:etd-02112004-115540
Authors: Yoler, Yusuf Amon
Year: 1954
DOI: 10.7907/853G-K555
The feasibility of using a shock tube for quantitative investigations of hypersonic flow phenomena at temperatures simulating free flight conditions is studied theoretically and experimentally. In the theoretical part, various aspects of the hypersonic shock tube problem are treated in logical order. Methods of producing high Mach numbers, limitations on the test section Mach number, methods of generating strong shock waves, flows with variable specific heats and dissociation, types of problems amenable to study with the hypersonic shock tube are discussed.
To verify and supplement some of the theoretical results, a shock tube of a somewhat unconventional design has been built. The bulk of the experimental investigations undertaken to date have dealt with pressure studies using piezoelectric gages, and schlieren studies of the flow, The results obtained so far with flow Mach numbers in excess of six, stagnation temperatures up to 9000[degrees]R and stagnation pressures up to 200 psi, have not only contributed to a much greater understanding of this relatively new field of application of the shock tube, but have indicated a well defined course along Which future investigations will continue.https://thesis.library.caltech.edu/id/eprint/604Viscous Effects on Static Pressure Distribution for a Slender Cone at a Nominal Mach Number of 5.8
https://resolver.caltech.edu/CaltechETD:etd-11242003-104401
Authors: Baldwin, Lawrence Cloyd
Year: 1955
DOI: 10.7907/XNB2-P865
An experimental investigation was conducted in the GALCIT Hypersonic Wind Tunnel, Leg No. 1, to determine the static pressure distribution on a cone with 5[degrees] semivertex angle at a nominal Mach number of 5.8.
This investigation was concerned with the effect of hypersonic boundary layer-shock wave interaction on the pressure at the cone surface. Pressure distributions were measured for three values of Reynolds numbers per inch and a comparison was made with theoretically calculated pressure distributions.
The influence of viscosity in hypersonic flow was demonstrated by an induced pressure rise of approximately 45% above theoretical inviscid pressure for the lowest Reynolds number tested.https://thesis.library.caltech.edu/id/eprint/4648An Experimental Investigation of Pressure Gradients Due to Temperature Gradients in Small Diameter Tubes
https://resolver.caltech.edu/CaltechETD:etd-12162003-155556
Authors: Howard, Weston Morgan
Year: 1955
DOI: 10.7907/SQF1-5002
NOTE: Text or symbols not renderable in plain ASCII are indicated by [...]. Abstraact is included in .pdf document.
Results of an experimental investigation of pressure gradients due to axial temperature gradients in small diameter tubes are presented. The tests, which covered the region of Knudsen numbers (based on tube inside radius) of 0.01 to 6, indicate good correlation with theory.
It is of value to note that this correlation was obtained by using [Delta]T equal to the temperature difference between the hot and cold ends of the tubes and T[subscript ave] equal to the average of these two temperatures. In contrast, theory would dictate obtaining the temperature variation along the length of the tube and applying the formulas to small incremental [Delta]T's, then summing to get the total effect. Therefore, for normal laboratory conditions where pressure gradient corrections are to be computed, it is sufficient to record only the temperatures at the hot and cold ends rather than having to obtain a number of temperature readings along the tube.
In order to apply pressure corrections easily and rapidly, a system of correction curves is given. To simplify the procedure, the tube cold end temperature was assumed to be 80[degrees]F, and the correction curves drawn accordingly. However, for different laboratory conditions a similar system of curves could be drawn and used.https://thesis.library.caltech.edu/id/eprint/5014An Experimental Investigation of Leading Edge Shock Wave-Boundary Layer Interaction at Hypersonic Speeds
https://resolver.caltech.edu/CaltechETD:etd-03252004-153340
Authors: Kendall, James Madison
Year: 1956
DOI: 10.7907/6XVK-B331
<p>The boundary layer on a slender body tends to be very thick at hypersonic speeds. It interacts with the external flow by producing larger flow deflections near the leading edge than those due to the body alone. The increased shock strength affects the boundary layer growth. The flow around the boundary layer gives rise to an induced pressure with a negative gradient which thins the boundary layer and increases the skin friction with respect to the zero pressure gradient value.</p>
<p>Experiments on a flat plate with a sharp leading edge (Reₜ < 100) have been performed in the GALCIT 5 x 5 inch Mach 5.8 hypersonic wind tunnel. The induced pressure was measured by means of orifices in the plate surface. Profiles of Mach number, velocity, mass flow, pressure, and momentum deficiency were calculated from impact pressure surveys normal to the plate surface made at various distances from the leading edge.</p>
<p>The results are as follows: (1) The induced pressures are 25 per cent higher than the weak interaction theory. (2) The boundary layer and the external flow are distinctly separate for Reₓ as low as 6000. (3) The shock wave location is in good agreement with that predicted by the Friedrichs theory for a body shape equivalent to the observed boundary layer displacement thickness. (4) Expansion waves reflected from the shock are weak. (5) The average skin friction coefficient tends toward and nearly matches the zero pressure gradient value downstream, but increases to approximately twice that value as the leading edge is approached.</p>https://thesis.library.caltech.edu/id/eprint/1112