Armand Sigalla

Bio: Armand Sigalla is an academic researcher from Boeing Commercial Airplanes. The author has contributed to research in topics: Aerodynamics & Drag. The author has an hindex of 3, co-authored 3 publications receiving 49 citations.

Real Flow Limitations in Supersonic Airplane Design

Abstract: angle of attack for this example i s found to be approximately 15 degrees. Experience indicates that the airstream normally Experimental studies, including pressure measurements, will not be able to flow around the leading edge at this large force measurements and flow visualization techniques, have angle of attack without flow separation. This i s particularly shown that predicted aerodynamic performance levels of supertrue for the thin airfoils that are characteristic of supersonic sonic wings can be achieved only when the flow remains wing designs. This leading.edge flow separation completely attached over the entire wing surface. alters the character o f the flow pattern over the wing. The nature of the breakdown of potential flow on supersonic wings is discussed and illustrated with experimental f low visualization pictures and wind-tunnel data. Various types of flow breakdown are examined. Simplified flow analogies that explain these flow phenomena are developed. Practical procedures that ensure design for attached flow at prescribed con. ditions are described. Flow analogies are used to explore the impact of various airplane design parameters on the breakdown of attached flow. 1.0 Introduction The design of efficient, very highly swept supersonic wings is one of the more difficult problems in aeronautics. These highly swept wings are of interest because they have the potential, according to theory, o f having relatively low drag a t super. sonic lifting conditions. LWell known supersonic wing theory' indicates that the leading edge of a wing must be a t an angle of sweepback greater than the angle weak shockwaves make with the free stream a t corresponding Mach numbers to achieve low drag a t lifting conditions. Sweepback angles of 70 to 75 degrees are necessary for Mach numbers in the range of 2.0 to 3.0 Theoretical predictions indicate that an airplane with a wing of such high sweep would have an advantage of approximately 15 to 20 percent in liftldrag ratio when compared to an airplane having a much lower sweepback angle (for example, 50 degrees). When i t was first attempted to substantiate these very encouraging predictions with wind-tunnel models, it was found that the experimental results did not confirm them a t all. Subsequent examinations revealed that the low drag predicted by theory was not achieved because the flow pattern around the wings, implicit in theory, did not occur in practice. Viscosity, which normally has a relatively small effect on the overall f low over wings a t normal cruise lift conditions, had a rather substantial effect on these highly swept wings. Consider as an example a wing a t Mach 3.0 and a t an angle of attack of 4 degrees-typical supersonic conditions. With the wing swept 75 degrees to achieve low drag, the Mach number component normal t o the leading edge i s 0.78. Hence near the wing leading edge, a recognized subsonic flow condition i s produced. The leading edge flow is governed by the angle normal to the leading edge. Using simple sweep theory. the normal !d Leadingedge flow separation i s only one of the reasons why the predicted low drag levels o f highly swept wings could not be obtained. The flow over the wing, which i s a t a relatively low pressure, must adjust t o freestream pressure through a shock wave at the trailing edge. If the theoretical f low requires too large a pressure rise, trailing-edge separation occurs. Again the flow pattern postulated by theory cannot occur and the theoretical drags cannot be achieved. Similar problems can occur on other partsof sucha highly swept wing. The establishment o f a f low consistent with theoretical low drag is, therefore, contingent on the response of the boundary layer to potentially severe conditions a l l over the wing. The development and behavior of highly swept wing boundary layers under complicated three-dimensional flow conditions is not amerable to theoretical calculations. Necessary wing design limitations cannot be defined strictly on the basis of analytical studies, and therefore had t o be developed from experimental test programs. This paper presents a comprehensive set o f design conditions that can be used to define efficient, highly swept super. sonic wings. I f these conditions are applied as constraints to theoretical calculations, the flow pattern resulting from analysis would not have a very large effect on the wing boundary layers and the theoretical flow, and drag, could be expected to be obtained in practice. The results presented in this paper are based on work that began in the la te 1950s and was carried through the U.S. SST program until cancellation of the program in 1971. The object of the work was to develop methods for the design of efficient supersonic wings. More recently. interest in the design of such wings has been renewed both for eventual commercial' and military3 applications. For the latter case, not only does the designer require low drag a t cruising conditions, but he also requires a reasonable flow at higher l i f t coefficients associated with military maneuvers. A review of design methods to accomplish this is therefore timely and appropriate, and forms the subject of this paper. In Section 2 the basic characteristics of supersonic wing planforms are discussed, pointing out the advantages of highly swept wings in supersonic flow. This i s followed by a review of experimental results illustrating the basic flow problems of highly swept wings. The potential effects of warping the sur. face of such wings, (e.g., camber and twist) are discussed in

Aerodynamics of powerplant installation on supersonic aircraft.

Abstract: Simple ideas derived from theoretical studies are utilized to explain the action of an engine nacelle upon a wing in supersonic flow. The engine installation is shown to influence both the wave drag due to thickness and the drag due to lift. Available theoretical procedures make it possible to estimate the forces on a wing due to the combined action of several nacelles, and results of theoretical calculations are compared with supersonic wind-tunnel-test data. The effects of the relative location of nacelles and the effects of nacelle shape and size on the lift, drag, and pitch characteristics of wind-tunnel models are presented and discussed. Experimental results are used to demonstrate that the camber and twist of the wing should be designed by taking into account the effects of the flowfield from the nacelles. It is indicated that designing for low drag requires consideration of the lift distribution on the airplane and a knowledge of viscous effects.

Real flow limitations in supersonic airplane design

Abstract: angle of attack for this example i s found to be approximately 15 degrees. Experience indicates that the airstream normally Experimental studies, including pressure measurements, will not be able to flow around the leading edge at this large force measurements and flow visualization techniques, have angle of attack without flow separation. This i s particularly shown that predicted aerodynamic performance levels of supertrue for the thin airfoils that are characteristic of supersonic sonic wings can be achieved only when the flow remains wing designs. This leading.edge flow separation completely attached over the entire wing surface. alters the character o f the flow pattern over the wing. The nature of the breakdown of potential flow on supersonic wings is discussed and illustrated with experimental f low visualization pictures and wind-tunnel data. Various types of flow breakdown are examined. Simplified flow analogies that explain these flow phenomena are developed. Practical procedures that ensure design for attached flow at prescribed con. ditions are described. Flow analogies are used to explore the impact of various airplane design parameters on the breakdown of attached flow. 1.0 Introduction The design of efficient, very highly swept supersonic wings is one of the more difficult problems in aeronautics. These highly swept wings are of interest because they have the potential, according to theory, o f having relatively low drag a t super. sonic lifting conditions. LWell known supersonic wing theory' indicates that the leading edge of a wing must be a t an angle of sweepback greater than the angle weak shockwaves make with the free stream a t corresponding Mach numbers to achieve low drag a t lifting conditions. Sweepback angles of 70 to 75 degrees are necessary for Mach numbers in the range of 2.0 to 3.0 Theoretical predictions indicate that an airplane with a wing of such high sweep would have an advantage of approximately 15 to 20 percent in liftldrag ratio when compared to an airplane having a much lower sweepback angle (for example, 50 degrees). When i t was first attempted to substantiate these very encouraging predictions with wind-tunnel models, it was found that the experimental results did not confirm them a t all. Subsequent examinations revealed that the low drag predicted by theory was not achieved because the flow pattern around the wings, implicit in theory, did not occur in practice. Viscosity, which normally has a relatively small effect on the overall f low over wings a t normal cruise lift conditions, had a rather substantial effect on these highly swept wings. Consider as an example a wing a t Mach 3.0 and a t an angle of attack of 4 degrees-typical supersonic conditions. With the wing swept 75 degrees to achieve low drag, the Mach number component normal t o the leading edge i s 0.78. Hence near the wing leading edge, a recognized subsonic flow condition i s produced. The leading edge flow is governed by the angle normal to the leading edge. Using simple sweep theory. the normal !d Leadingedge flow separation i s only one of the reasons why the predicted low drag levels o f highly swept wings could not be obtained. The flow over the wing, which i s a t a relatively low pressure, must adjust t o freestream pressure through a shock wave at the trailing edge. If the theoretical f low requires too large a pressure rise, trailing-edge separation occurs. Again the flow pattern postulated by theory cannot occur and the theoretical drags cannot be achieved. Similar problems can occur on other partsof sucha highly swept wing. The establishment o f a f low consistent with theoretical low drag is, therefore, contingent on the response of the boundary layer to potentially severe conditions a l l over the wing. The development and behavior of highly swept wing boundary layers under complicated three-dimensional flow conditions is not amerable to theoretical calculations. Necessary wing design limitations cannot be defined strictly on the basis of analytical studies, and therefore had t o be developed from experimental test programs. This paper presents a comprehensive set o f design conditions that can be used to define efficient, highly swept super. sonic wings. I f these conditions are applied as constraints to theoretical calculations, the flow pattern resulting from analysis would not have a very large effect on the wing boundary layers and the theoretical flow, and drag, could be expected to be obtained in practice. The results presented in this paper are based on work that began in the la te 1950s and was carried through the U.S. SST program until cancellation of the program in 1971. The object of the work was to develop methods for the design of efficient supersonic wings. More recently. interest in the design of such wings has been renewed both for eventual commercial' and military3 applications. For the latter case, not only does the designer require low drag a t cruising conditions, but he also requires a reasonable flow at higher l i f t coefficients associated with military maneuvers. A review of design methods to accomplish this is therefore timely and appropriate, and forms the subject of this paper. In Section 2 the basic characteristics of supersonic wing planforms are discussed, pointing out the advantages of highly swept wings in supersonic flow. This i s followed by a review of experimental results illustrating the basic flow problems of highly swept wings. The potential effects of warping the sur. face of such wings, (e.g., camber and twist) are discussed in

Leeside flows over delta wings at supersonic speeds

Abstract: An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

Aircraft Drag Reduction-A Review

Abstract: The paper summarizes the state of the art in aeronautical drag reduction across the speed range for the conventional drag components of viscous drag, drag due to lift and wave drag. It also describes several emerging drag-reduction approaches that are either active or reactive/interactive and the drag reduction potentially available from synergistic combinations of advanced configuration aerodynamics, viscous drag-reduction approaches, revolutionary structural concepts and propulsion integration.

Supersonic aircraft drag reduction

Abstract: This paper reviews aerodynamic drag reduction for friction, wave and vortex drag associated with supersonic cruise aircraft and suggests approaches and research directions. Suction laminar flow control may also enable improved low-speed, high-lift systems, improved lift-to-drag ratio for subsonic cruise, reduced parasitic viscous drag for favorable interference wave drag reduction approaches, and turbulent skin friction reduction via slot injection. Flow separation control at cruise proffers opportunities for increased leading-edge thrust, increased lift increment from upper surface, increased fuselage lift/camber for wave drag-due-to-lift reduction, improved performance of various favorable interference wave drag reduction schemes, as well as possibly better low-speed, high-lift systems and wing cruise performance.

Real Flow Limitations in Supersonic Airplane Design

Abstract: angle of attack for this example i s found to be approximately 15 degrees. Experience indicates that the airstream normally Experimental studies, including pressure measurements, will not be able to flow around the leading edge at this large force measurements and flow visualization techniques, have angle of attack without flow separation. This i s particularly shown that predicted aerodynamic performance levels of supertrue for the thin airfoils that are characteristic of supersonic sonic wings can be achieved only when the flow remains wing designs. This leading.edge flow separation completely attached over the entire wing surface. alters the character o f the flow pattern over the wing. The nature of the breakdown of potential flow on supersonic wings is discussed and illustrated with experimental f low visualization pictures and wind-tunnel data. Various types of flow breakdown are examined. Simplified flow analogies that explain these flow phenomena are developed. Practical procedures that ensure design for attached flow at prescribed con. ditions are described. Flow analogies are used to explore the impact of various airplane design parameters on the breakdown of attached flow. 1.0 Introduction The design of efficient, very highly swept supersonic wings is one of the more difficult problems in aeronautics. These highly swept wings are of interest because they have the potential, according to theory, o f having relatively low drag a t super. sonic lifting conditions. LWell known supersonic wing theory' indicates that the leading edge of a wing must be a t an angle of sweepback greater than the angle weak shockwaves make with the free stream a t corresponding Mach numbers to achieve low drag a t lifting conditions. Sweepback angles of 70 to 75 degrees are necessary for Mach numbers in the range of 2.0 to 3.0 Theoretical predictions indicate that an airplane with a wing of such high sweep would have an advantage of approximately 15 to 20 percent in liftldrag ratio when compared to an airplane having a much lower sweepback angle (for example, 50 degrees). When i t was first attempted to substantiate these very encouraging predictions with wind-tunnel models, it was found that the experimental results did not confirm them a t all. Subsequent examinations revealed that the low drag predicted by theory was not achieved because the flow pattern around the wings, implicit in theory, did not occur in practice. Viscosity, which normally has a relatively small effect on the overall f low over wings a t normal cruise lift conditions, had a rather substantial effect on these highly swept wings. Consider as an example a wing a t Mach 3.0 and a t an angle of attack of 4 degrees-typical supersonic conditions. With the wing swept 75 degrees to achieve low drag, the Mach number component normal t o the leading edge i s 0.78. Hence near the wing leading edge, a recognized subsonic flow condition i s produced. The leading edge flow is governed by the angle normal to the leading edge. Using simple sweep theory. the normal !d Leadingedge flow separation i s only one of the reasons why the predicted low drag levels o f highly swept wings could not be obtained. The flow over the wing, which i s a t a relatively low pressure, must adjust t o freestream pressure through a shock wave at the trailing edge. If the theoretical f low requires too large a pressure rise, trailing-edge separation occurs. Again the flow pattern postulated by theory cannot occur and the theoretical drags cannot be achieved. Similar problems can occur on other partsof sucha highly swept wing. The establishment o f a f low consistent with theoretical low drag is, therefore, contingent on the response of the boundary layer to potentially severe conditions a l l over the wing. The development and behavior of highly swept wing boundary layers under complicated three-dimensional flow conditions is not amerable to theoretical calculations. Necessary wing design limitations cannot be defined strictly on the basis of analytical studies, and therefore had t o be developed from experimental test programs. This paper presents a comprehensive set o f design conditions that can be used to define efficient, highly swept super. sonic wings. I f these conditions are applied as constraints to theoretical calculations, the flow pattern resulting from analysis would not have a very large effect on the wing boundary layers and the theoretical flow, and drag, could be expected to be obtained in practice. The results presented in this paper are based on work that began in the la te 1950s and was carried through the U.S. SST program until cancellation of the program in 1971. The object of the work was to develop methods for the design of efficient supersonic wings. More recently. interest in the design of such wings has been renewed both for eventual commercial' and military3 applications. For the latter case, not only does the designer require low drag a t cruising conditions, but he also requires a reasonable flow at higher l i f t coefficients associated with military maneuvers. A review of design methods to accomplish this is therefore timely and appropriate, and forms the subject of this paper. In Section 2 the basic characteristics of supersonic wing planforms are discussed, pointing out the advantages of highly swept wings in supersonic flow. This i s followed by a review of experimental results illustrating the basic flow problems of highly swept wings. The potential effects of warping the sur. face of such wings, (e.g., camber and twist) are discussed in

A system for aerodynamic design and analysis of supersonic aircraft. Part 2: User's manual

Abstract: The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics were included in the system to display or edit input and to permit monitoring and readout of program results.