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Research Reports and Memoranda

Research Reports and Memoranda F.193. Stability Diagram for Laminar Boundary Layer Flow. R. Timman, J. Λ. Zaat and Th. J. Research Reports and Memoranda Burgerhout. October 1956. Stability diagrams are given for a one-parameter family of boundary layer velocity profiles. The Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical method of calculation applied starts from the asymp­ Research Council, Reports and Technical Memoranda of the United States National Advisory Com­ totic behaviour of the inviscid differential equation of disturbance. This method makes use of only one mittee for Aeronautics and publications of other similar Research Bodies as issued. solution in the complete domain of integration. TECHNICAL NOTES CANADA TECHNICAL REPORTS 24. An Experimental Investition of a Supersonic NATIONAL RESEARCH COUNCIL F.208. The Influence of Non-Stationary Stability Two-Dimensional Perforated Inlet at a Nominal Free 100 Sussex Drive, Ottawa 2 Derivatives on the Snaking Motion at High Subsonic Stream Mach Number of 2·5 . By J . P. С Clark. Speed. By J . IJff. November 1957. AERONAUTICAL REPORTS November 1958. Some lateral stability calculations with two degrees LR-259. An Experimental Study of Rotating Stall A two-dimensional supersonic diffuser in the form of freedom, for flight at high subsonic speed, have in a Two-stage Axial Compressor. By R. F. Meyer. of a reversed de Laval nozzle has been tested at a been performed in order to study the influence of August 1959. nominal free stream Mach number of 2· 5. The re­ non-stationary stability derivatives on the damping The stalling behaviour of an Orenda CT-100 ex­ sulting bow shock was induced to move into the intake of the snaking motion. perimental two-stage axial compressor was studied at and eventually past the throat by perforating the the constant operating speed of 3,000 r.p.m. Rotating diffuser walls. Windows in the side walls enabled REPORTS AND TRANSACTIONS stall was observed to commence at a mass flow co­ visual studies of the shock swallowing process to be efficient of 0·435 and as the mass flow was reduced made and a series of 19 static pressure taps in the XXII. Boundary Value Problems in Lifting Surface several different patterns of rotating stall were ob­ diffuser walls provided pressure distributions as a Theory. By E . Van Spiegel. March 1959. served. In two ranges of operation 'steady' patterns function of perforation and exit areas. Stable opera­ A lifting surface of circular planform in steady and consisting of three rotating stalls were observed, in tion of the diffuser was demonstrated with the shock unsteady incompressible flow is considered. The two other ranges of operation rotating stall was in any position from detached to downstream of the problems are formulated as boundary value problems observed, but the stall patterns were continually throat. A maximum pressure recovery of 87· 5 per for the Laplace equation. Appropriate orthogonal changing, while for very low mass flow operation the cent was obtained at the throat with an associated co-ordinates are introduced. Solutions of the Laplace flow over the whole of the annulus appeared to be mass flow loss of 21 per cent. There were indications equation arc found by separation of variables and stalled. Reversed flow within the stalls was found to however that proper allowance for boundary layer Green's function of the second kind can be found. be an important feature of rotating stall. growth at the throat could reduce the loss value to The complete solution of each of the physical prob­ about 12· 5 per cent. lems can be written as the sum of the regular acceler­ ation potential and a singular solution which is LR-260. Approximate Method for Determining the singular only along the leading edge of the wing. Potential Flow about an Arbitrary Aerofoil Section in a ITALY Two-Dimensional Finite Stream with Particular Refer­ U.S.A. ence to Large Stream Deflections. By D . G. Gould. MINISTERO DIFESA-AERONAUTICA August 1959. Centro Consultivo Studi e Ricerche NATIONAL AERONAUTICS AND SPACE Rome The flow about and on the surface of an aerofoil ADMINISTRATION section in a finite stream and that about and on the MONOGRAPHS Government Printing Office, Washington 25, D.C. surface of the same aerofoil section in an infinite sym­ 5. Fuselage structures of non-circular section, TECHNICAL TRANSLATIONS metrical cascade arc compared. It is shown that, for (Fusoliere a sezione non circolare.) By P . Santini, F.1 . The Characteristics of Hydrogen and Water most cases of practical interest, the differences arc June 1959. as Working Gases for Reactorheatcd Rocket Motors. small for stream deflexion angles up to 90 degrees. The problem of the distribution of stresses and dis­ By Sänger-Bredt. (Trans, from Astronautica Acta, Thus, the approximate flow about an arbitrary aero­ placement in a fuselage structure of non-circular sec­ v. 3, No. 4, 1957.) foil section in a two-dimensional finite stream may be tion, with rigid ribs, is considered. The Balance obtained by using the methods existing for the deter­ Hydrogen and water vapour have been investigated Method is used. The complete solution of the problem mination of the flow about an arbitrary section in a as to their applicability as working fluids with an is given, with some typical results which provide an cascade. The procedure, as applied to the finite stream energy higher by several orders of magnitude than idea of the effect of non-circularity. approximation, is given using the interference method has hitherto been available with steam boilers. of cascade theory. Enthalpy-entropy diagrams are computed for either gas, with complete regard to dissociation, as well as JAPAN ionization, over the pressure range between 10+1 to MECHANICAL ENGINEERING REPORTS 10-5 atm., and within the temperature range between AERONAUTICAL RESEARCH INSTITUTE 500 and 10,000 deg. K. Exhaust velocities, possible for MP-15. The Solution of an Eigenvalue Problem UNIVERSITY OF TOKYO water, or water vapour, are derived and compared. (Rate of Evaporation and Decomposition of a Droplet Komabu, Meguro-ku, Tokyo Heat transfer to the combustion chamber walls is of an Ozone-Oxygen Mixture) by Means of Analogue REPORTS examined. Three methods of heating the working fluids and Digital Computers. By R. Sandri. August 1959. are investigated: fission reactor, arcs, and fusion 347. The Stress Distribution in a Swept-back Box- The rate of evaporation and decomposition of a reactors. beam under Torsional and Bending Loads. By K. droplet of a 50-50 mole per cent mixture of liquid Ikeda and M . Sunakawa. July 1959. ozone and oxygen in a hot oxygen atmosphere at 20 atmospheres pressure was computed for different sizes F.2. On the Temperature Distribution Behind Cylin­ An approximate solution is presented to obtain the of the droplet radius. This is an eigenvalue problem ders in a Flow. By J . Ackeret. (Trans. from Mitteil­ static stress distributions in a swept-back box-beam which could be solved with an accuracy of about ungen 21, Institut für Aerodynamik, E.T.H., 1954.) having asymmetrical section. The box-beam is as­ 2 per cent on the PACE analogue computer. As an sumed to consist of four concentrated flanges of A simple theoretical explanation of the aerodynamic accuracy check, the one-dimensional problem (rate of different cross-sectional area and four thin walls of cooling often found near the base and in the wake of evaporation of a large drop) was solved also on the different thickness, which connect the flanges. The bluff bodies is presented. The theory is based on com­ Bendix digital computer. The result differed by ribs are assumed to be perfectly stiff, except the one pressible potential flow and makes use of the energy 2·2 per cent from that obtained on the analogue nearest to the root. increment associated with the unsteady velocity- computer. potential term of the energy equation in calculating the temperature at a point in unsteady flow for the general case. The case of two-dimensional flow past a NETHERLANDS INSTITUTE OF AEROPHYSICS cylinder is examined in detail. The calculated tem­ NATIONAAL LUCHTVAARTLABORATORIUM University of Toronto perature distribution around a cylinder, assuming (N.L.L.) incompressible flow, is compared with a distribution REPORTS Slatcrwcg 145, Amsterdam found experimentally at M=0·35. The comparison 57. An Experimental Investigation of the Noise TECHNICAL NOTES tends to verify the theoretical explanation. Generated by the Turbulent Flow around a Rotating Cylinder. By L . N. Wilson. April 1959. F.192. Calculation of Aerodynamic Forces on Slowly Oscillating Rectangular Wings in Subsonic F-3. Piecemeal Solutions in the Programming of The near and far field noise from the turbulent Flow. By A. I. van de Vooren and E. M. de Jager. Optimal Flight Trajectories. (Soluzioni discontinue boundary layer developed on a rigid wall rotating October 1956. nei problemi di volo ottimo.) By P. Cicala. October cylinder has been studied ; both smooth and artificially 1959. (Translation from Accademia delle Scienze di roughened surfaces were employed. The mean square A method is presented for the calculation of the Torino, June 1956.) pressure followed aU law in the near field, and ap­ aerodynamic forces on a slowly oscillatingacrofoil. The proximated to a U06 law, as did the acoustic power, in method is essentially a lifting surface theory which The general equations that govern the shape of the the far field. The U06 law suggests that the dominant takes into account the unsteady effects due to the optimal flight paths that may be executed in a vertical radiators are of a dipole type. Drag calculations from wake. Results are given as scries containing terms of plane arc written out in a form which is especially hot wire measurements provided an estimate of the increasing powers in reduced frequency. It is shown suited to study of the practical situation wherein the efficiency of the far field radiation, resulting in an that if the axis of rotation is ahead of the aerofoil, the complete trajectory is made up out of piecemeal sec­ efficiency about ten times that for the quadruple radia­ aerodynamic damping is much less than it would be tions of diverse character. As a rule, real flight paths tion from a jet at a Mach number of 0·228. Even so, according to quasi-steady theory. Instability is possible will be found to consist of such spliced together por­ the measured noise power from the boundary layer if the aspect ratio is larger than a certain value. This tions of trajectory segments, along which special local over a rigid wall appears to be relatively small in any limiting aspect ratio decreases with increasing Mach rules of flight will be in force in order to meet a series practical application to the noise in aircraft. number. of variant conditions. February 1960 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 32 (2): 1 – Feb 1, 1960
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Emerald Publishing
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Copyright © Emerald Group Publishing Limited
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0002-2667
DOI
10.1108/eb033213
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Abstract

F.193. Stability Diagram for Laminar Boundary Layer Flow. R. Timman, J. Λ. Zaat and Th. J. Research Reports and Memoranda Burgerhout. October 1956. Stability diagrams are given for a one-parameter family of boundary layer velocity profiles. The Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical method of calculation applied starts from the asymp­ Research Council, Reports and Technical Memoranda of the United States National Advisory Com­ totic behaviour of the inviscid differential equation of disturbance. This method makes use of only one mittee for Aeronautics and publications of other similar Research Bodies as issued. solution in the complete domain of integration. TECHNICAL NOTES CANADA TECHNICAL REPORTS 24. An Experimental Investition of a Supersonic NATIONAL RESEARCH COUNCIL F.208. The Influence of Non-Stationary Stability Two-Dimensional Perforated Inlet at a Nominal Free 100 Sussex Drive, Ottawa 2 Derivatives on the Snaking Motion at High Subsonic Stream Mach Number of 2·5 . By J . P. С Clark. Speed. By J . IJff. November 1957. AERONAUTICAL REPORTS November 1958. Some lateral stability calculations with two degrees LR-259. An Experimental Study of Rotating Stall A two-dimensional supersonic diffuser in the form of freedom, for flight at high subsonic speed, have in a Two-stage Axial Compressor. By R. F. Meyer. of a reversed de Laval nozzle has been tested at a been performed in order to study the influence of August 1959. nominal free stream Mach number of 2· 5. The re­ non-stationary stability derivatives on the damping The stalling behaviour of an Orenda CT-100 ex­ sulting bow shock was induced to move into the intake of the snaking motion. perimental two-stage axial compressor was studied at and eventually past the throat by perforating the the constant operating speed of 3,000 r.p.m. Rotating diffuser walls. Windows in the side walls enabled REPORTS AND TRANSACTIONS stall was observed to commence at a mass flow co­ visual studies of the shock swallowing process to be efficient of 0·435 and as the mass flow was reduced made and a series of 19 static pressure taps in the XXII. Boundary Value Problems in Lifting Surface several different patterns of rotating stall were ob­ diffuser walls provided pressure distributions as a Theory. By E . Van Spiegel. March 1959. served. In two ranges of operation 'steady' patterns function of perforation and exit areas. Stable opera­ A lifting surface of circular planform in steady and consisting of three rotating stalls were observed, in tion of the diffuser was demonstrated with the shock unsteady incompressible flow is considered. The two other ranges of operation rotating stall was in any position from detached to downstream of the problems are formulated as boundary value problems observed, but the stall patterns were continually throat. A maximum pressure recovery of 87· 5 per for the Laplace equation. Appropriate orthogonal changing, while for very low mass flow operation the cent was obtained at the throat with an associated co-ordinates are introduced. Solutions of the Laplace flow over the whole of the annulus appeared to be mass flow loss of 21 per cent. There were indications equation arc found by separation of variables and stalled. Reversed flow within the stalls was found to however that proper allowance for boundary layer Green's function of the second kind can be found. be an important feature of rotating stall. growth at the throat could reduce the loss value to The complete solution of each of the physical prob­ about 12· 5 per cent. lems can be written as the sum of the regular acceler­ ation potential and a singular solution which is LR-260. Approximate Method for Determining the singular only along the leading edge of the wing. Potential Flow about an Arbitrary Aerofoil Section in a ITALY Two-Dimensional Finite Stream with Particular Refer­ U.S.A. ence to Large Stream Deflections. By D . G. Gould. MINISTERO DIFESA-AERONAUTICA August 1959. Centro Consultivo Studi e Ricerche NATIONAL AERONAUTICS AND SPACE Rome The flow about and on the surface of an aerofoil ADMINISTRATION section in a finite stream and that about and on the MONOGRAPHS Government Printing Office, Washington 25, D.C. surface of the same aerofoil section in an infinite sym­ 5. Fuselage structures of non-circular section, TECHNICAL TRANSLATIONS metrical cascade arc compared. It is shown that, for (Fusoliere a sezione non circolare.) By P . Santini, F.1 . The Characteristics of Hydrogen and Water most cases of practical interest, the differences arc June 1959. as Working Gases for Reactorheatcd Rocket Motors. small for stream deflexion angles up to 90 degrees. The problem of the distribution of stresses and dis­ By Sänger-Bredt. (Trans, from Astronautica Acta, Thus, the approximate flow about an arbitrary aero­ placement in a fuselage structure of non-circular sec­ v. 3, No. 4, 1957.) foil section in a two-dimensional finite stream may be tion, with rigid ribs, is considered. The Balance obtained by using the methods existing for the deter­ Hydrogen and water vapour have been investigated Method is used. The complete solution of the problem mination of the flow about an arbitrary section in a as to their applicability as working fluids with an is given, with some typical results which provide an cascade. The procedure, as applied to the finite stream energy higher by several orders of magnitude than idea of the effect of non-circularity. approximation, is given using the interference method has hitherto been available with steam boilers. of cascade theory. Enthalpy-entropy diagrams are computed for either gas, with complete regard to dissociation, as well as JAPAN ionization, over the pressure range between 10+1 to MECHANICAL ENGINEERING REPORTS 10-5 atm., and within the temperature range between AERONAUTICAL RESEARCH INSTITUTE 500 and 10,000 deg. K. Exhaust velocities, possible for MP-15. The Solution of an Eigenvalue Problem UNIVERSITY OF TOKYO water, or water vapour, are derived and compared. (Rate of Evaporation and Decomposition of a Droplet Komabu, Meguro-ku, Tokyo Heat transfer to the combustion chamber walls is of an Ozone-Oxygen Mixture) by Means of Analogue REPORTS examined. Three methods of heating the working fluids and Digital Computers. By R. Sandri. August 1959. are investigated: fission reactor, arcs, and fusion 347. The Stress Distribution in a Swept-back Box- The rate of evaporation and decomposition of a reactors. beam under Torsional and Bending Loads. By K. droplet of a 50-50 mole per cent mixture of liquid Ikeda and M . Sunakawa. July 1959. ozone and oxygen in a hot oxygen atmosphere at 20 atmospheres pressure was computed for different sizes F.2. On the Temperature Distribution Behind Cylin­ An approximate solution is presented to obtain the of the droplet radius. This is an eigenvalue problem ders in a Flow. By J . Ackeret. (Trans. from Mitteil­ static stress distributions in a swept-back box-beam which could be solved with an accuracy of about ungen 21, Institut für Aerodynamik, E.T.H., 1954.) having asymmetrical section. The box-beam is as­ 2 per cent on the PACE analogue computer. As an sumed to consist of four concentrated flanges of A simple theoretical explanation of the aerodynamic accuracy check, the one-dimensional problem (rate of different cross-sectional area and four thin walls of cooling often found near the base and in the wake of evaporation of a large drop) was solved also on the different thickness, which connect the flanges. The bluff bodies is presented. The theory is based on com­ Bendix digital computer. The result differed by ribs are assumed to be perfectly stiff, except the one pressible potential flow and makes use of the energy 2·2 per cent from that obtained on the analogue nearest to the root. increment associated with the unsteady velocity- computer. potential term of the energy equation in calculating the temperature at a point in unsteady flow for the general case. The case of two-dimensional flow past a NETHERLANDS INSTITUTE OF AEROPHYSICS cylinder is examined in detail. The calculated tem­ NATIONAAL LUCHTVAARTLABORATORIUM University of Toronto perature distribution around a cylinder, assuming (N.L.L.) incompressible flow, is compared with a distribution REPORTS Slatcrwcg 145, Amsterdam found experimentally at M=0·35. The comparison 57. An Experimental Investigation of the Noise TECHNICAL NOTES tends to verify the theoretical explanation. Generated by the Turbulent Flow around a Rotating Cylinder. By L . N. Wilson. April 1959. F.192. Calculation of Aerodynamic Forces on Slowly Oscillating Rectangular Wings in Subsonic F-3. Piecemeal Solutions in the Programming of The near and far field noise from the turbulent Flow. By A. I. van de Vooren and E. M. de Jager. Optimal Flight Trajectories. (Soluzioni discontinue boundary layer developed on a rigid wall rotating October 1956. nei problemi di volo ottimo.) By P. Cicala. October cylinder has been studied ; both smooth and artificially 1959. (Translation from Accademia delle Scienze di roughened surfaces were employed. The mean square A method is presented for the calculation of the Torino, June 1956.) pressure followed aU law in the near field, and ap­ aerodynamic forces on a slowly oscillatingacrofoil. The proximated to a U06 law, as did the acoustic power, in method is essentially a lifting surface theory which The general equations that govern the shape of the the far field. The U06 law suggests that the dominant takes into account the unsteady effects due to the optimal flight paths that may be executed in a vertical radiators are of a dipole type. Drag calculations from wake. Results are given as scries containing terms of plane arc written out in a form which is especially hot wire measurements provided an estimate of the increasing powers in reduced frequency. It is shown suited to study of the practical situation wherein the efficiency of the far field radiation, resulting in an that if the axis of rotation is ahead of the aerofoil, the complete trajectory is made up out of piecemeal sec­ efficiency about ten times that for the quadruple radia­ aerodynamic damping is much less than it would be tions of diverse character. As a rule, real flight paths tion from a jet at a Mach number of 0·228. Even so, according to quasi-steady theory. Instability is possible will be found to consist of such spliced together por­ the measured noise power from the boundary layer if the aspect ratio is larger than a certain value. This tions of trajectory segments, along which special local over a rigid wall appears to be relatively small in any limiting aspect ratio decreases with increasing Mach rules of flight will be in force in order to meet a series practical application to the noise in aircraft. number. of variant conditions. February 1960

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Aircraft Engineering and Aerospace Technology – Emerald Publishing

Published: Feb 1, 1960

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