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Research Reports and Memoranda

Research Reports and Memoranda A I R C R A F T E N G I N E E R I N G 73 September 1966 3387. The Turbulent Boundary Layer with Suction or Injection. By T. J. Black and A. J. Sarnecki. Research Reports and Memoranda October 1958. This report considers the turbulent boundary layer with distributed suction or injection applied normally Under this heading are published regularly abstracts of Reports and Memoranda of the Aeronautical through the surface. A bilogarithmic law of the wall is established analogous to the logarithmic law for Research Council and publications of other similar Research Bodies as issued. impervious surfaces. Coles' wake hypothesis is ex­ tended to the transpiration layer and verified experi­ mentally. The wall boundary condition is discussed of various multi-shock and isentropic compression GREAT BRITAIN briefly and possible effects of surface irregularity arc flows, and of the lateral slenderness of related caret examined. Finally the overall picture of the turbulent AERONAUTICAL RESEARCH COUNCIL bodies. For bodies producing isentropic compression, transpiration layer is discussed. It is considered that spanwise pressure distributions arc shown to depend H.M. Stationery Office, London this report provides an acceptable framework for the on anhedral and planform and it is noted that such evolution of a complete theory. REPORTS AND MEMORANDA bodies should be of use in research on three-dimen­ sional boundary layers. It is shown that multi-wave and isentropic compression surfaces may be used as 3388. The effect of Forward Speed on the Inlet Flow intakes (or isentropic bodies as nozzles), and that 3380. Some Tests on High-Reaction Compressor Distribution and Performance of a Lifting Fan In­ these can be used as components in the design of con­ Blanding. By R. C.Turner and R. A. Barrows. January stalled in a Wing. By N. Gregory, W. G. Raymer and figurations, beneath which lifting and propulsive 1963. (8s. 6d.) Edna M. Love. June 1962. (16s. 6d.) flows can be wholly two-dimensional. Various models Two sets of blading, designed for 120 per cent are proposed for inclusion in current wind-tunnel Yawmeter traverses show that flow maldistributions reaction, and for flow coefficients of 0·667 and 1·0 programmes. due to forward speed can be removed by a deep duct respectively, have been tested in the N.G.T.E. low- or an inlet cascade. Non-uniformity of exit static speed compressor No. 106. The performances were pressure does not affect the flow distribution. Since compared with those of corresponding sets of 50-per- forward speed reduces the pressure rise required, 3384. The Development of a Nozzle for Absolute cent-reaction blading and with predictions based on stalling of the fan blades appears not to be a great Airflow Measurement by Pitot-Static Traverse. simple theoretical methods. The high-reaction blad­ danger except at very high speeds, or when a deflected By J . C. Ascough. May 1963. (20s.) ings both gave efficiencies which were lower by exit cascade is fitted. amounts broadly in line with the predictions. The work- A special nozzle has been made with associated done factors were also appreciably lower. The lower- ducting and instruments in order to provide an abso­ flow-coefficient blading showed a significant advan­ lute measure of airflow. The equipment could form a 3390. Measurements at Subsonic and Supersonic tage in pressure rise over its 50-per-cent-reaction portable self-contained assembly, enabling calibra­ Speeds of the Longitudinal and Lateral Stability of a counterpart, with about the same surge flow, while the tions to be made of meters installed in test rigs. In Slender Cambered Ogee Wing including the Effects of a high-flow blading was deficient in pressure rise but had order to give an absolute measurement, the nozzle has Fin, Canopy Nose and Trailing-Edge Controls. By D. a considerably wider surge margin. Both bladings been designed to generate an idealized flow amenable Isaacs. September 1963. might thus be of value in specialised situations where both to theoretical prediction and to experimental efficiency is not a prime consideration. The results show that for longitudinal stability at survey. Firstly, the nozzle produces a uniform velocity M=0· 3 and Ce=0·45, the centre of gravity of an profile across the mainstream, which can easily be traversed for pitot and static pressure to a high order of actual aircraft could be located only forward of 66 per accuracy. Secondly, the boundary layer is fairly thin cent c . The centre of pressure of the wing with basic 3381. Measurements of Aerodynamic Derivatives on in the traverse plane, thus minimizing the effect of nose and no fin is at 71 per cent c at the cruise atti­ a Wing with a Series of Tip Bodies. By P . R. Guyett. variations within it on nozzle C . The uncertainty of tude, M=2· 2 and C =0·075, so that the camber used D L March 1963. (20s.) an airflow measurement in a steady flow is estimated is insufficient to trim the wing. Measured values of the as ±0·17 per cent due to known random errors. By drag increments due to control deflection show fair Oscillatory aerodynamic lift and pitching-moment far the biggest error is that resulting from circum­ agreement with linear-theory estimates. The control derivatives have been measured in a wind tunnel at ferential variation of the boundary-layer profiles. It is effectiveness can be predicted with fair accuracy. The low subsonic speeds on the following wing and wing- felt that much of the error from the boundary layer canopy nose is slightly de-stabilizing in yaw, and it body combinations: (i) tapered wing of aspect ratio could be removed with the manufacture of a new has a drag penalty which is probably larger than could 1·3, mid-chord line unswept, with a sharp leading be tolerated (30 per cent of basic wing wave drag). nozzle. An extensive subsidiary experimental pro­ edge; (ii) above wing fitted with a nacelle in three At supersonic speeds slender-body theory is generally gramme has been carried out to solve the problem of chordwise positions at the tip; (iii) wing at (i) fitted inadequate for predicting the lateral derivatives of the the accurate measurement of static pressure. The with a tip body representing a tank or store, and effect of static-hole size was found to agree with wing. The fin effectiveness can be estimated with good tested with and without a fin at the rear of the body. Shaw's correlation. accuracy. The measured derivatives for the wing alone are in good agreement with values calculated using lifting- surface theory, and arc in reasonable agreement with values calculated using a semi-empirical method. 3391. On the Shear Flexibility of Twisted Panels. By 3385. Transonic Tunnel Tests on a 6 per cent Thick, Derivatives for the wing with nacelle in the central S. Kelsey and D. F . Pilkington. April 1964. Warped 55 deg. Sweptback-Wing Model. By A. B. position have been calculated using a semi-empirical Haines and J . C. M. Jones. September 1960 (17s.) A simple theoretical analysis is given for the addi­ method which makes an allowance for aerodynamic tional shear flexibility of a twisted panel caused by Tests have been made in the A.R.A. 9 ft. by 8 ft. interference effects, and comparison shows that the bending of the panel under a shear loading. Results of transonic tunnel on a model having a 6 per cent thick, values are reasonably accurate. some experiments on initially twisted sandwich panels 55 deg. sweptback wing with a warp distribution de­ show good agreement with the theory. signed to give a constant spanwise C —distribution and a triangular chordwise load at C =0·15, M— 1·2. The wing-body junction was designed according to 3382. An Experiment on Aerodynamic Nozzles at 3393. Boundary-Layer Drag of Bi-Convex Wing supersonic area rule for this Mach number. Some M = 2. By J . Reid. November 1962. (15s.) Sections with Heat Transfer at Supersonic Speeds. By incremental drags are also given for a particular fin- In an aerodynamic nozzle a convergent primary R. E. Luxton and A. D. Young. March 1964. tail unit but these are not necessarily representative of nozzle is housed within an afterbody which is sur­ a full-scale aircraft since the fin-tail unit was designed Calculations are made for a bi-convex wing of 5 per rounded by a shroud. Part of the external flow is specifically to suit the requirements of a subsequent cent thickness and a flat plate at Mach numbers of l·5 , captured thereby and compressed by the under-ex­ free-flight test of the model. Finally, the report in­ 2 ·5 and 5·0, and Reynolds numbers of 106, 107 and 108 panded primary jet. At the design point this system is cludes some data showing the downwash at the tail for a range of heat-transfer conditions. For the wing a not as efficient as a convergent-divergent nozzle but its position both for this model and for a corresponding wide range of transition positions are covered. It is off-design performance should be better. In the model having a symmetrical untwisted wing. shown that the effect of rearward movement of experiment axi-symmetric models were used. The transition is reduced by increase of Mach number, by shrouds were cylindrical externally and either cylin­ reduction of Reynolds number and by increase in the drical or convergent-divergent internally. With each ratio of wall temperature to recovery temperature. model the thrust and afterbody pressure distributions 3386. Bending Flutter of Unstalled Cascade Blades These results are explained in terms of the relative were measured at jet pressure ratios up to 20 with a at Finite Deflection. By D . S . Whitehead. October 1962 sensitivities of the laminar and turbulent boundary free-stream Mach number of 2·0. Component drags This report gives a method of calculating the condi­ layers to these parameters. It is shown that in general were estimated indirectly from these measurements. tions under which a cascade of unstalled blades will cooling of the surface causes an increase in drag for all For comparison, the thrust of unshrouded convergent flutter in their bending mode. The blades are assumed transition points except those very close to the trailing and convergent-divergent nozzles was measured under to be flat plates and the flow is assumed to be two- edge. the same conditions. dimensional. It is found that, when there is deflection of the steady flow through the cascade, bending flutter can occur in which there is a phase difference between 3394. Notes on the Analysis of Stability in Ac­ the motion of one blade and its neighbours. The flutter celerated Motion. By S. B. Gates and A. W. Thorpe. 3383. On Lifting Bodies which contain Two-Dimen­ boundaries for a range of cascades have been calcu­ September 1964. sional Supersonic Flows. By L. H. Townsend. August lated on a digital computer and arc presented. For 1963. (18s.) turbine cascades these boundaries are not far from Linear differential equations whose coefficients arc typical operating conditions in gas turbines and this functions of the independent variable occur in the This paper is primarily concerned with the design may be why blades in the low pressure end of steam discussion of the motion following a small disturbance principles of three-dimensional surfaces which can in a specified accelerated motion. An indirect attack on turbines have been found to require lacing wires. produce two-dimensional, centred compression waves this problem is made by seeking to establish upper Comparisons arc made with earlier theories due to and are derived from the Nonweiler wing. A study is bounds to the solution of second-order equations. made of the efficiencies (i.e. of the pressure recoveries) Sohngen and Shioiri. http://www.deepdyve.com/assets/images/DeepDyve-Logo-lg.png Aircraft Engineering and Aerospace Technology Emerald Publishing

Research Reports and Memoranda

Aircraft Engineering and Aerospace Technology , Volume 38 (9): 1 – Sep 1, 1966
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Publisher
Emerald Publishing
Copyright
Copyright © Emerald Group Publishing Limited
ISSN
0002-2667
DOI
10.1108/eb034191
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Abstract

A I R C R A F T E N G I N E E R I N G 73 September 1966 3387. The Turbulent Boundary Layer with Suction or Injection. By T. J. Black and A. J. Sarnecki. Research Reports and Memoranda October 1958. This report considers the turbulent boundary layer with distributed suction or injection applied normally Under this heading are published regularly abstracts of Reports and Memoranda of the Aeronautical through the surface. A bilogarithmic law of the wall is established analogous to the logarithmic law for Research Council and publications of other similar Research Bodies as issued. impervious surfaces. Coles' wake hypothesis is ex­ tended to the transpiration layer and verified experi­ mentally. The wall boundary condition is discussed of various multi-shock and isentropic compression GREAT BRITAIN briefly and possible effects of surface irregularity arc flows, and of the lateral slenderness of related caret examined. Finally the overall picture of the turbulent AERONAUTICAL RESEARCH COUNCIL bodies. For bodies producing isentropic compression, transpiration layer is discussed. It is considered that spanwise pressure distributions arc shown to depend H.M. Stationery Office, London this report provides an acceptable framework for the on anhedral and planform and it is noted that such evolution of a complete theory. REPORTS AND MEMORANDA bodies should be of use in research on three-dimen­ sional boundary layers. It is shown that multi-wave and isentropic compression surfaces may be used as 3388. The effect of Forward Speed on the Inlet Flow intakes (or isentropic bodies as nozzles), and that 3380. Some Tests on High-Reaction Compressor Distribution and Performance of a Lifting Fan In­ these can be used as components in the design of con­ Blanding. By R. C.Turner and R. A. Barrows. January stalled in a Wing. By N. Gregory, W. G. Raymer and figurations, beneath which lifting and propulsive 1963. (8s. 6d.) Edna M. Love. June 1962. (16s. 6d.) flows can be wholly two-dimensional. Various models Two sets of blading, designed for 120 per cent are proposed for inclusion in current wind-tunnel Yawmeter traverses show that flow maldistributions reaction, and for flow coefficients of 0·667 and 1·0 programmes. due to forward speed can be removed by a deep duct respectively, have been tested in the N.G.T.E. low- or an inlet cascade. Non-uniformity of exit static speed compressor No. 106. The performances were pressure does not affect the flow distribution. Since compared with those of corresponding sets of 50-per- forward speed reduces the pressure rise required, 3384. The Development of a Nozzle for Absolute cent-reaction blading and with predictions based on stalling of the fan blades appears not to be a great Airflow Measurement by Pitot-Static Traverse. simple theoretical methods. The high-reaction blad­ danger except at very high speeds, or when a deflected By J . C. Ascough. May 1963. (20s.) ings both gave efficiencies which were lower by exit cascade is fitted. amounts broadly in line with the predictions. The work- A special nozzle has been made with associated done factors were also appreciably lower. The lower- ducting and instruments in order to provide an abso­ flow-coefficient blading showed a significant advan­ lute measure of airflow. The equipment could form a 3390. Measurements at Subsonic and Supersonic tage in pressure rise over its 50-per-cent-reaction portable self-contained assembly, enabling calibra­ Speeds of the Longitudinal and Lateral Stability of a counterpart, with about the same surge flow, while the tions to be made of meters installed in test rigs. In Slender Cambered Ogee Wing including the Effects of a high-flow blading was deficient in pressure rise but had order to give an absolute measurement, the nozzle has Fin, Canopy Nose and Trailing-Edge Controls. By D. a considerably wider surge margin. Both bladings been designed to generate an idealized flow amenable Isaacs. September 1963. might thus be of value in specialised situations where both to theoretical prediction and to experimental efficiency is not a prime consideration. The results show that for longitudinal stability at survey. Firstly, the nozzle produces a uniform velocity M=0· 3 and Ce=0·45, the centre of gravity of an profile across the mainstream, which can easily be traversed for pitot and static pressure to a high order of actual aircraft could be located only forward of 66 per accuracy. Secondly, the boundary layer is fairly thin cent c . The centre of pressure of the wing with basic 3381. Measurements of Aerodynamic Derivatives on in the traverse plane, thus minimizing the effect of nose and no fin is at 71 per cent c at the cruise atti­ a Wing with a Series of Tip Bodies. By P . R. Guyett. variations within it on nozzle C . The uncertainty of tude, M=2· 2 and C =0·075, so that the camber used D L March 1963. (20s.) an airflow measurement in a steady flow is estimated is insufficient to trim the wing. Measured values of the as ±0·17 per cent due to known random errors. By drag increments due to control deflection show fair Oscillatory aerodynamic lift and pitching-moment far the biggest error is that resulting from circum­ agreement with linear-theory estimates. The control derivatives have been measured in a wind tunnel at ferential variation of the boundary-layer profiles. It is effectiveness can be predicted with fair accuracy. The low subsonic speeds on the following wing and wing- felt that much of the error from the boundary layer canopy nose is slightly de-stabilizing in yaw, and it body combinations: (i) tapered wing of aspect ratio could be removed with the manufacture of a new has a drag penalty which is probably larger than could 1·3, mid-chord line unswept, with a sharp leading be tolerated (30 per cent of basic wing wave drag). nozzle. An extensive subsidiary experimental pro­ edge; (ii) above wing fitted with a nacelle in three At supersonic speeds slender-body theory is generally gramme has been carried out to solve the problem of chordwise positions at the tip; (iii) wing at (i) fitted inadequate for predicting the lateral derivatives of the the accurate measurement of static pressure. The with a tip body representing a tank or store, and effect of static-hole size was found to agree with wing. The fin effectiveness can be estimated with good tested with and without a fin at the rear of the body. Shaw's correlation. accuracy. The measured derivatives for the wing alone are in good agreement with values calculated using lifting- surface theory, and arc in reasonable agreement with values calculated using a semi-empirical method. 3391. On the Shear Flexibility of Twisted Panels. By 3385. Transonic Tunnel Tests on a 6 per cent Thick, Derivatives for the wing with nacelle in the central S. Kelsey and D. F . Pilkington. April 1964. Warped 55 deg. Sweptback-Wing Model. By A. B. position have been calculated using a semi-empirical Haines and J . C. M. Jones. September 1960 (17s.) A simple theoretical analysis is given for the addi­ method which makes an allowance for aerodynamic tional shear flexibility of a twisted panel caused by Tests have been made in the A.R.A. 9 ft. by 8 ft. interference effects, and comparison shows that the bending of the panel under a shear loading. Results of transonic tunnel on a model having a 6 per cent thick, values are reasonably accurate. some experiments on initially twisted sandwich panels 55 deg. sweptback wing with a warp distribution de­ show good agreement with the theory. signed to give a constant spanwise C —distribution and a triangular chordwise load at C =0·15, M— 1·2. The wing-body junction was designed according to 3382. An Experiment on Aerodynamic Nozzles at 3393. Boundary-Layer Drag of Bi-Convex Wing supersonic area rule for this Mach number. Some M = 2. By J . Reid. November 1962. (15s.) Sections with Heat Transfer at Supersonic Speeds. By incremental drags are also given for a particular fin- In an aerodynamic nozzle a convergent primary R. E. Luxton and A. D. Young. March 1964. tail unit but these are not necessarily representative of nozzle is housed within an afterbody which is sur­ a full-scale aircraft since the fin-tail unit was designed Calculations are made for a bi-convex wing of 5 per rounded by a shroud. Part of the external flow is specifically to suit the requirements of a subsequent cent thickness and a flat plate at Mach numbers of l·5 , captured thereby and compressed by the under-ex­ free-flight test of the model. Finally, the report in­ 2 ·5 and 5·0, and Reynolds numbers of 106, 107 and 108 panded primary jet. At the design point this system is cludes some data showing the downwash at the tail for a range of heat-transfer conditions. For the wing a not as efficient as a convergent-divergent nozzle but its position both for this model and for a corresponding wide range of transition positions are covered. It is off-design performance should be better. In the model having a symmetrical untwisted wing. shown that the effect of rearward movement of experiment axi-symmetric models were used. The transition is reduced by increase of Mach number, by shrouds were cylindrical externally and either cylin­ reduction of Reynolds number and by increase in the drical or convergent-divergent internally. With each ratio of wall temperature to recovery temperature. model the thrust and afterbody pressure distributions 3386. Bending Flutter of Unstalled Cascade Blades These results are explained in terms of the relative were measured at jet pressure ratios up to 20 with a at Finite Deflection. By D . S . Whitehead. October 1962 sensitivities of the laminar and turbulent boundary free-stream Mach number of 2·0. Component drags This report gives a method of calculating the condi­ layers to these parameters. It is shown that in general were estimated indirectly from these measurements. tions under which a cascade of unstalled blades will cooling of the surface causes an increase in drag for all For comparison, the thrust of unshrouded convergent flutter in their bending mode. The blades are assumed transition points except those very close to the trailing and convergent-divergent nozzles was measured under to be flat plates and the flow is assumed to be two- edge. the same conditions. dimensional. It is found that, when there is deflection of the steady flow through the cascade, bending flutter can occur in which there is a phase difference between 3394. Notes on the Analysis of Stability in Ac­ the motion of one blade and its neighbours. The flutter celerated Motion. By S. B. Gates and A. W. Thorpe. 3383. On Lifting Bodies which contain Two-Dimen­ boundaries for a range of cascades have been calcu­ September 1964. sional Supersonic Flows. By L. H. Townsend. August lated on a digital computer and arc presented. For 1963. (18s.) turbine cascades these boundaries are not far from Linear differential equations whose coefficients arc typical operating conditions in gas turbines and this functions of the independent variable occur in the This paper is primarily concerned with the design may be why blades in the low pressure end of steam discussion of the motion following a small disturbance principles of three-dimensional surfaces which can in a specified accelerated motion. An indirect attack on turbines have been found to require lacing wires. produce two-dimensional, centred compression waves this problem is made by seeking to establish upper Comparisons arc made with earlier theories due to and are derived from the Nonweiler wing. A study is bounds to the solution of second-order equations. made of the efficiencies (i.e. of the pressure recoveries) Sohngen and Shioiri.

Journal

Aircraft Engineering and Aerospace Technology – Emerald Publishing

Published: Sep 1, 1966

There are no references for this article.