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Research Reports and Memoranda

Research Reports and Memoranda LR-409. Tests on a Tip Turbine Volute with Circular Cross-Sections and a Gooseneck Outlet. By R. J. Kind. October 1964. Model tests were made on a volute designed by a Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical method that takes into account wall friction, a free vortex flow pattern, and compressibility. The tests Research Council, Technical Reports and Translations of the United States National Aeronautics and indicated that a free vortex flow pattern docs prevail Space Administration and publications of other similar Research Bodies as issued. over most of the volute, with some deviation at both the beginning and end of the scroll. Because of a GREAT BRITAIN ference on wings of finite span in a rectangular wind secondary flow in the volute, the effect of friction tunnel with closed side walls and porous-slotted AERONAUTICAL RESEARCH COUNCIL differs from that anticipated in the design method. floor and roof. A resume of previous results con­ H. M. Stationery Office, London Near the re-entry region the outcoming flow shows a cerning two-dimensional aerofoils and small-span greater deviation than elsewhere from ideal be­ REPORTS AND MEMORANDA wings is included for completeness. haviour. 3338. Low-speed wind-tunnel measurements on a 3397. Measurements of the Direct Pitching Oscilla­ thin sharp-edged Delta wing with 70° leading-edge MECHANICAL ENGINEERING REPORT tion Derivatives for Three Cropped Delta and Three sweep, with particular reference to the position of MS-111. The Relationship of Fatigue Endurance of Arrowhead Planforms at Subsonic and Transonic leading-edge-Vortex breakdown. By J. A. Lawford 2024-T3 Aluminium Alloy to Atmospheric Humidity. speeds. By C.J . W. Miles and K. B. Bridgman. and A. R. Beauchamp. November 1961. (6s.) By J . A. Dunsby and W. Wiebe. September 1964. August 1962. (12s.) The position of the breakdown of the tightly rolled A brief review is made of the present state of know­ Measurements of direct pitching oscillation deri­ leading-edge vortex was observed using a smoke ledge on the effects of atmospheric humidity on the vative coefficients at subsonic and transonic speeds technique. The breakdown point moved forward with fatigue strength of metals. Experiments are described have been made on three cropped delta wings of increasing incidence, and reached a point above the that were performed to determine how these humidity aspect 3·0, 2·0 and 1·35 and three arrowhead plan- trailing edge at an incidence of 32 deg. It moved for­ effects vary over a range of stress level and, also, to forms having varying degrees of sweepback. Results ward to transverse planes through 0·5 and 0·28 of determine whether relative or absolute humidity has are presented for measurements about two model the centre-line chord at 34 deg. and 37 deg. incidence the controlling influence for specimens of 2024-T3 axes and for two amplitudes of oscillation. Comparison respectively. The root-mean-square intensity and the aluminium alloy. with theory shows reasonable agreement. low-frequency component of pressure fluctuations both began to rise rapidly at approximately the same incidence (31 deg.) at four widely spaced points on the 3398. The Change in Shock-Tunnel Tailoring Mach TEST REPORTS wing. Number due to Driver Gas Mixtures of Helium and MET-452. A Smoke Generator for Use in Wind Nitrogen. By L. Pennelegion and P. J. Gough. Tunnels. By H . S. Fowler. August 1964. October 1963. (8s.) 3345. Dynamics of the Deformable Aeroplane: In order to overcome disadvantages in existing Part 1, the Equations of Motion; Part 2, The Study Comparison is made of the pressure recovery and methods, a new wind tunnel smoke system has been of the Trim State and Longitudinal Stability of the available testing time of a shock tunnel using room- investigated. The system consists simply of a device Slender Integrated Aeroplane Configuration. By R. D. temperature nitrogen, helium or a mixture of these as for mixing jets of anhydrous ammonia and sulphur Milne. September 1962. (£1 5s.) the driver gas. It is shown that tailoring can be dioxide gases. It is instantly controllable, produces obtained from M =1·0 to 3·4, resulting in an An integrated analytical treatment is presented which S dense white smoke without clogging, is easily portable, improvement in pressure recovery and testing time. deals with the equilibrium and stability of the flexible and poses no fire risk. Provided simple precautions aeroplane in flight. The analysis embodies those are taken, there is no danger of toxic concentrations 3399. Notes and Comments on Some Particular methods currently employed to investigate the building up and the substances used are non-corrosive, behaviour of the flexible aeroplane stemming on the features of Wind-Tunnel Design Following Measure­ readily available, and cheap. one hand from the stability theory of the rigid aero­ ments of Flow Distribution, Wall Pressures, etc. in the plane and on the other from conventional aerolastic N.P.L. 7 ft. (Low Turbulence) Wind Tunnel. By MET-453. Pressure Loss Tests on a Tip-Turbine studies. The integrated treatment serves to clarify C. Salter and W. G. Rayner. October 1963. (9s. 6d.) Volute Model having a Gooseneck Outlet Passage the regions of application of these restricted methods. Fitted With Nozzle Blades. By M . J . Hamer. Septem­ Various measurements are recorded and some of In Part I the equations of motion for a flexible aero­ ber 1964. the features (good and otherwise) of the design are plane are developed in as general a manner as possible. discussed. The conclusions have relevance both to the Pressure loss tests were run on a model of a tip- In Part II the general analysis is applied to a detailed modification of existing wind tunnels and to the turbine volute duet fitted with a complete set of study of the equilibrium and stability of the slender, design of new ones. nozzle blades. The overall pressure loss was found to integrated aeroplane configuration. be 12·9 9 percent of the supply total pressure (gauge). CANADA This was subdivided into 7·2 per cent nozzle loss and 3349. A contribution to the theory of aircraft 5·7 per cent volute loss. The overall pressure loss NATIONAL RESEARCH COUNCIL OF CANADA response in rolling manoeuvres including inertia cross- pattern, the outlet velocities and angles were all more 100 Sussex Drive, Ottawa, 2. coupling effects. By H. H. B. M. Thomas and P. uniform around the discharge plane than with the Price. April 1960. (£1 5s.) AERONAUTICAL REPORTS bladeless volute. The problem of calculating the response of an air­ LR-389. A Theoretical Study of the Inviscid Hyper­ craft in rolling manoeuvres when the mass distribution sonic Flow about a Conical Flat-top Wind-body INSTITUTE FOR AEROSPACE STUDIES of the aircraft is such that the inertia terms in the Combination. By P . Mandl. January 1964. University of Toronto. equations of motion effect a cross-coupling of the By combining the theory of linearized characteris­ TECHNICAL NOTES usual lateral and longitudinal motions is considered. tics with the hypersonic small disturbance approxi­ Solutions are outlined to two formulations of this 70. An Experimental Investigation into the Shape of mation, explicit expressions are derived for the shock problem: (i) Response to a given applied aileron and Thrust Augmenting Surfaces in Conjunction with shape, the flow field, the surface pressure distribution (ii) Response corresponding to a specified time history Coanda-Deflected Jet Sheets (Part I). By C. D. Hope- and the aerodynamic forces for a delta wing and half of rate of roll. Detailed calculations are made only Gill. July 1964. cone combination travelling at hypersonic speeds for the first of these, and the results compare favour­ and small incidence. The theory is applied to several It is known that by means of additional surfaces, the ably with digital-computer solutions. configurations for which experimental data on surface thrust of a jet or jet sheet can be enhanced (thrust Possible simplifications to the method of calcula­ pressures and stream-line inclination are available. augmentation). This especially applies to Coanda- tion are discussed. Theoretical and experimental values agree quite deflected jet sheets because of the inherent stronger closely on the body surface, but in the wing-body entrainment into curved flow surfaces. The flow from 3368. Three-Dimensional Disturbances in a Two- junction and on the wing the theory predicts pressures a two-dimensional subsonic nozzle was deflected by Dimensional Incompressible Turbulent Boundary Layer. that are too low. These discrepancies are believed to quadrants. A composite thrust-augmenting surface By H. Fernholz. October 1962. (6s. 6d.) be due to complicated viscous interaction phenomena was added, and the effect of its shape on thrust aug­ Measurements in turbulent boundary layers have occurring in the junction, which are not taken into mentation was studied at various nozzle pressure shown large transverse variations of skin friction. It account by the theory. ratios and radii of the quadrants. This investigation appears that these are associated with three-dimen­ yielded a maximum thrust augmentation of 1·21 for sional disturbances originating in the transition region. several optimum configurations, which was governed LR-407. Gas Turbine Cycle Calculations: Design- The disturbances cannot be explained by irregularities primarily by the relative magnitude and direction of Point Performance of Turbojet and Turbofan Engines. in the damping screens and cannot be eliminated by the momentum of the secondary (entrained) flow in By M. S. Chappell, E. P. Cockshutt and C. R. fixing transition. relation to the primary (nozzle) flow momentum. Sharp. October 1964. Thrust augmentation decreased with increasing An accurate and comprehensive method for cal­ nozzle pressure ratio but was independent of quadrant 3392. On the Axial Compression of Long Slightly culating the design-point performance of turbojet radius. Curved Panels. By G. G. Pope. October 1963. (12s.) and turbofan engines is described in this report. A finite-deflection analysis is given of the buckling Analyses of the component processes that comprise a 86. On the Lateral Instabilities of Aircraft Due to of long, slightly curved panels in compression parallel cycle are developed in detail from fundamental Parametric Excitation. By M. Masak. January 1965. to the generators, with sides either clamped or simply- thermodynamic relations to the actual equations supported, and with various combinations of boundary programmed for the computer. The structure and Lateral instabilities of aircraft due to parametric conditions in the middle surface. excitation are studied. A theoretical discussion in­ coding of the Fortran II computer programme volving asymptotic approximations to the state are outlined, and complete source listings are included 3395. Upwash Interference on Wings of Finite transition matrix and a Liapunov stability analysis as an appendix. Sufficiently detailed instructions Span in a Rectangular Wind Tunnel with Closed Side regarding input-output techniques and programme are presented first. This is followed by an analogue Walls and Porous-Slotted Floor and Roof. By D. R. execution are given to enable cycle studies to be computer study to observe the actual phenomenon. Holder. November 1963. (4s. 6d.) carried out by those without intimate knowledge of Both high-aspect-ration swept-wing and slender delta- the computer or the programme. wing jet transports are considered. An exact solution is obtained for the upwash inter­ June 1965 193 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 37 (6): 1 – Jun 1, 1965

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Publisher
Emerald Publishing
Copyright
Copyright © Emerald Group Publishing Limited
ISSN
0002-2667
DOI
10.1108/eb034031
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Abstract

LR-409. Tests on a Tip Turbine Volute with Circular Cross-Sections and a Gooseneck Outlet. By R. J. Kind. October 1964. Model tests were made on a volute designed by a Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical method that takes into account wall friction, a free vortex flow pattern, and compressibility. The tests Research Council, Technical Reports and Translations of the United States National Aeronautics and indicated that a free vortex flow pattern docs prevail Space Administration and publications of other similar Research Bodies as issued. over most of the volute, with some deviation at both the beginning and end of the scroll. Because of a GREAT BRITAIN ference on wings of finite span in a rectangular wind secondary flow in the volute, the effect of friction tunnel with closed side walls and porous-slotted AERONAUTICAL RESEARCH COUNCIL differs from that anticipated in the design method. floor and roof. A resume of previous results con­ H. M. Stationery Office, London Near the re-entry region the outcoming flow shows a cerning two-dimensional aerofoils and small-span greater deviation than elsewhere from ideal be­ REPORTS AND MEMORANDA wings is included for completeness. haviour. 3338. Low-speed wind-tunnel measurements on a 3397. Measurements of the Direct Pitching Oscilla­ thin sharp-edged Delta wing with 70° leading-edge MECHANICAL ENGINEERING REPORT tion Derivatives for Three Cropped Delta and Three sweep, with particular reference to the position of MS-111. The Relationship of Fatigue Endurance of Arrowhead Planforms at Subsonic and Transonic leading-edge-Vortex breakdown. By J. A. Lawford 2024-T3 Aluminium Alloy to Atmospheric Humidity. speeds. By C.J . W. Miles and K. B. Bridgman. and A. R. Beauchamp. November 1961. (6s.) By J . A. Dunsby and W. Wiebe. September 1964. August 1962. (12s.) The position of the breakdown of the tightly rolled A brief review is made of the present state of know­ Measurements of direct pitching oscillation deri­ leading-edge vortex was observed using a smoke ledge on the effects of atmospheric humidity on the vative coefficients at subsonic and transonic speeds technique. The breakdown point moved forward with fatigue strength of metals. Experiments are described have been made on three cropped delta wings of increasing incidence, and reached a point above the that were performed to determine how these humidity aspect 3·0, 2·0 and 1·35 and three arrowhead plan- trailing edge at an incidence of 32 deg. It moved for­ effects vary over a range of stress level and, also, to forms having varying degrees of sweepback. Results ward to transverse planes through 0·5 and 0·28 of determine whether relative or absolute humidity has are presented for measurements about two model the centre-line chord at 34 deg. and 37 deg. incidence the controlling influence for specimens of 2024-T3 axes and for two amplitudes of oscillation. Comparison respectively. The root-mean-square intensity and the aluminium alloy. with theory shows reasonable agreement. low-frequency component of pressure fluctuations both began to rise rapidly at approximately the same incidence (31 deg.) at four widely spaced points on the 3398. The Change in Shock-Tunnel Tailoring Mach TEST REPORTS wing. Number due to Driver Gas Mixtures of Helium and MET-452. A Smoke Generator for Use in Wind Nitrogen. By L. Pennelegion and P. J. Gough. Tunnels. By H . S. Fowler. August 1964. October 1963. (8s.) 3345. Dynamics of the Deformable Aeroplane: In order to overcome disadvantages in existing Part 1, the Equations of Motion; Part 2, The Study Comparison is made of the pressure recovery and methods, a new wind tunnel smoke system has been of the Trim State and Longitudinal Stability of the available testing time of a shock tunnel using room- investigated. The system consists simply of a device Slender Integrated Aeroplane Configuration. By R. D. temperature nitrogen, helium or a mixture of these as for mixing jets of anhydrous ammonia and sulphur Milne. September 1962. (£1 5s.) the driver gas. It is shown that tailoring can be dioxide gases. It is instantly controllable, produces obtained from M =1·0 to 3·4, resulting in an An integrated analytical treatment is presented which S dense white smoke without clogging, is easily portable, improvement in pressure recovery and testing time. deals with the equilibrium and stability of the flexible and poses no fire risk. Provided simple precautions aeroplane in flight. The analysis embodies those are taken, there is no danger of toxic concentrations 3399. Notes and Comments on Some Particular methods currently employed to investigate the building up and the substances used are non-corrosive, behaviour of the flexible aeroplane stemming on the features of Wind-Tunnel Design Following Measure­ readily available, and cheap. one hand from the stability theory of the rigid aero­ ments of Flow Distribution, Wall Pressures, etc. in the plane and on the other from conventional aerolastic N.P.L. 7 ft. (Low Turbulence) Wind Tunnel. By MET-453. Pressure Loss Tests on a Tip-Turbine studies. The integrated treatment serves to clarify C. Salter and W. G. Rayner. October 1963. (9s. 6d.) Volute Model having a Gooseneck Outlet Passage the regions of application of these restricted methods. Fitted With Nozzle Blades. By M . J . Hamer. Septem­ Various measurements are recorded and some of In Part I the equations of motion for a flexible aero­ ber 1964. the features (good and otherwise) of the design are plane are developed in as general a manner as possible. discussed. The conclusions have relevance both to the Pressure loss tests were run on a model of a tip- In Part II the general analysis is applied to a detailed modification of existing wind tunnels and to the turbine volute duet fitted with a complete set of study of the equilibrium and stability of the slender, design of new ones. nozzle blades. The overall pressure loss was found to integrated aeroplane configuration. be 12·9 9 percent of the supply total pressure (gauge). CANADA This was subdivided into 7·2 per cent nozzle loss and 3349. A contribution to the theory of aircraft 5·7 per cent volute loss. The overall pressure loss NATIONAL RESEARCH COUNCIL OF CANADA response in rolling manoeuvres including inertia cross- pattern, the outlet velocities and angles were all more 100 Sussex Drive, Ottawa, 2. coupling effects. By H. H. B. M. Thomas and P. uniform around the discharge plane than with the Price. April 1960. (£1 5s.) AERONAUTICAL REPORTS bladeless volute. The problem of calculating the response of an air­ LR-389. A Theoretical Study of the Inviscid Hyper­ craft in rolling manoeuvres when the mass distribution sonic Flow about a Conical Flat-top Wind-body INSTITUTE FOR AEROSPACE STUDIES of the aircraft is such that the inertia terms in the Combination. By P . Mandl. January 1964. University of Toronto. equations of motion effect a cross-coupling of the By combining the theory of linearized characteris­ TECHNICAL NOTES usual lateral and longitudinal motions is considered. tics with the hypersonic small disturbance approxi­ Solutions are outlined to two formulations of this 70. An Experimental Investigation into the Shape of mation, explicit expressions are derived for the shock problem: (i) Response to a given applied aileron and Thrust Augmenting Surfaces in Conjunction with shape, the flow field, the surface pressure distribution (ii) Response corresponding to a specified time history Coanda-Deflected Jet Sheets (Part I). By C. D. Hope- and the aerodynamic forces for a delta wing and half of rate of roll. Detailed calculations are made only Gill. July 1964. cone combination travelling at hypersonic speeds for the first of these, and the results compare favour­ and small incidence. The theory is applied to several It is known that by means of additional surfaces, the ably with digital-computer solutions. configurations for which experimental data on surface thrust of a jet or jet sheet can be enhanced (thrust Possible simplifications to the method of calcula­ pressures and stream-line inclination are available. augmentation). This especially applies to Coanda- tion are discussed. Theoretical and experimental values agree quite deflected jet sheets because of the inherent stronger closely on the body surface, but in the wing-body entrainment into curved flow surfaces. The flow from 3368. Three-Dimensional Disturbances in a Two- junction and on the wing the theory predicts pressures a two-dimensional subsonic nozzle was deflected by Dimensional Incompressible Turbulent Boundary Layer. that are too low. These discrepancies are believed to quadrants. A composite thrust-augmenting surface By H. Fernholz. October 1962. (6s. 6d.) be due to complicated viscous interaction phenomena was added, and the effect of its shape on thrust aug­ Measurements in turbulent boundary layers have occurring in the junction, which are not taken into mentation was studied at various nozzle pressure shown large transverse variations of skin friction. It account by the theory. ratios and radii of the quadrants. This investigation appears that these are associated with three-dimen­ yielded a maximum thrust augmentation of 1·21 for sional disturbances originating in the transition region. several optimum configurations, which was governed LR-407. Gas Turbine Cycle Calculations: Design- The disturbances cannot be explained by irregularities primarily by the relative magnitude and direction of Point Performance of Turbojet and Turbofan Engines. in the damping screens and cannot be eliminated by the momentum of the secondary (entrained) flow in By M. S. Chappell, E. P. Cockshutt and C. R. fixing transition. relation to the primary (nozzle) flow momentum. Sharp. October 1964. Thrust augmentation decreased with increasing An accurate and comprehensive method for cal­ nozzle pressure ratio but was independent of quadrant 3392. On the Axial Compression of Long Slightly culating the design-point performance of turbojet radius. Curved Panels. By G. G. Pope. October 1963. (12s.) and turbofan engines is described in this report. A finite-deflection analysis is given of the buckling Analyses of the component processes that comprise a 86. On the Lateral Instabilities of Aircraft Due to of long, slightly curved panels in compression parallel cycle are developed in detail from fundamental Parametric Excitation. By M. Masak. January 1965. to the generators, with sides either clamped or simply- thermodynamic relations to the actual equations supported, and with various combinations of boundary programmed for the computer. The structure and Lateral instabilities of aircraft due to parametric conditions in the middle surface. excitation are studied. A theoretical discussion in­ coding of the Fortran II computer programme volving asymptotic approximations to the state are outlined, and complete source listings are included 3395. Upwash Interference on Wings of Finite transition matrix and a Liapunov stability analysis as an appendix. Sufficiently detailed instructions Span in a Rectangular Wind Tunnel with Closed Side regarding input-output techniques and programme are presented first. This is followed by an analogue Walls and Porous-Slotted Floor and Roof. By D. R. execution are given to enable cycle studies to be computer study to observe the actual phenomenon. Holder. November 1963. (4s. 6d.) carried out by those without intimate knowledge of Both high-aspect-ration swept-wing and slender delta- the computer or the programme. wing jet transports are considered. An exact solution is obtained for the upwash inter­ June 1965 193

Journal

Aircraft Engineering and Aerospace TechnologyEmerald Publishing

Published: Jun 1, 1965

There are no references for this article.