JPH01500742A - Transonic wing design procedure - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed description of the invention] Transonic wing design procedure Technical field The present invention relates to an airplane wing design procedure. In particular, the present invention operates at transonic speeds. Concerning procedures for designing high-performance wings. Transonic means that the flight speed is the airflow around the airplane. The expansion (and acceleration) results in a supersonic airflow regime, thus mixing with the potential for shock waves. (Sub-supersonic) Refers to the flight speed range system below the speed of sound, which is an issue in airflow design. .

Background technology Typically, once the performance requirements are defined, a typical dimensional study is performed to determine the basic airfoil shape (sweep angle, taper, thickness and aspect ratio). Then, the two-dimensional airfoil The design is based on the designer's experience through repeated wind tunnel experiments or computer simulations. It has been developed using numerical and analytical techniques by Then this two-dimensional design Used as the basis for three-dimensional wing shape.

A three-dimensional wing developed from a two-dimensional wing cross section is carried out in a wind tunnel experiment or test flight. is almost always the case. The performance characteristics are consistent with the original expected performance based on the two-dimensional cross-sectional characteristics. There is no guarantee that it will be possible. Specifies the explanation of why a particular design does not meet performance requirements. Although various reasons are asserted, such explanations usually suggest that the wing is modified for good performance. It is not conducive to forming information about what can be formed. Instead of this, We repeated the design, experimentally changed it, and repeated wind tunnel experiments until we reached a perfect design. It is necessary to accomplish this. This trial-and-error approach is extremely expensive and is not always acceptable. Important considerations regarding the usefulness of conventional airfoil design methods It raises controversial points. Three-dimensional analysis or computer techniques can determine resistance levels. Although unpredictable, the requirements for estimating wing performance naturally evolve and Compromises in implementation for design.

Disclosure of invention According to the invention, the airfoil is constructed using the conventional steps of first defining the performance requirements and Designed using requirements to introduce dimensions.

According to the invention, the next step is to determine the aerodynamics of the wing with respect to the immediate position of the shock wave directly above. The purpose is to determine the sweepback angle A. This step is the next step, where the secondary The original Matsuha number Ml-o is the three-dimensional Matsuha number M, -0 and the cosine value co of the aerodynamic sweepback angle. The two-dimensional lift coefficient Ctz-0 is determined by multiplying by sΔ, and the two-dimensional lift coefficient Ctz-0 is It is calculated using the square cos2A of the cosine value of the receding angle.

Therefore, the shape of the two-dimensional airfoil is based on the two-dimensional Matsuha number and two-dimensional lift coefficient requirements. Determined by The shape of the wing in three dimensions is then determined by a series of lines along the chord line of the wing. at the position of the wing along the curved structure of the diagonal chord line perpendicular to the wing's local recession line. Defined by placing the airfoil.

According to the present invention, the above method can be applied to a plurality of positions along the span from the root of the blade to the tip of the blade. may be formed for each location.

Of course, the present invention uses the second-dimensional flow simulation technique to obtain the third-dimensional results. It also includes evaluating the performance of the original wing.

The design evaluation and refinement steps are repeated until the desired predictive performance characteristics are achieved. may be formed.

Brief description of the drawing In order to carry out the present invention easily and efficiently, it will be explained with reference to the accompanying drawings.

Figure 1 is a two-dimensional cross-sectional view of the airfoil in the x-Y plane, and Figure 2 is used for conventional design analysis. Figure 3 shows the concept of a conventional finite tapered wing. 4 is a conceptual diagram of a finitely tapered airfoil with a curved airfoil according to the present invention, and FIG. 5A- FIG. 5F is a series of backward tapers for each Matsuha number for the airfoil of FIG. 5G according to the present invention. Figure 6 shows the pressure distribution correlation using two-dimensional airfoil analysis combined with theory. The states obtained in Figures 5A to 5F for comparison with comparative values using regression theory. The table shown, FIG. 7, shows the improved prediction achieved using the backward taper theory according to the present invention. FIG. 3 is a diagram showing imaginary resistance.

BEST MODE FOR CARRYING OUT THE INVENTION Jet airplanes can operate at speeds slower than the speed of sound, but the airflow over the wings is slow. It is known that the supersonic range currently exists in the supersonic region.

Transonic flight speeds cause shock waves to exist on the wings. This shock wave increases in speed This increases drag and therefore limits aircraft performance. that Therefore, when designing a wing, it is important to avoid sharp large pressure changes that would indicate a severe shock wave. It is preferable to have a chord-like pressure contour in which the pressure gradually changes.

Conventional blade designs only consider secondary flow effects with simple adjustments for each blade sweep angle. was. According to such a conventional procedure, the blade is placed in one set in the X-Y plane of FIG. The design is based on an airfoil 10 that follows contour lines defined by shared coordinates. death However, the wing must be defined by contour lines in three-dimensional space. in the transonic region The wing used must have a taper and sweep . Chord length tapers are used for structural reasons to reduce wing weight. Recession The corners reduce wave drag by reducing the air velocity passing through the shock wave. ing.

In two-dimensional airfoil design, the cross-sectional pressure profile is adjusted so that the shock wave is as weak as possible. Attempts were made to contour. The cross-sectional thickness, lift force and Matsuha number are determined according to this ideal plan. It defines the limits of how close to However, this range formula has some strength The maximum range would be obtained if the plane flew at a speed that produced a shock wave of Usually indicates that

The designed two-dimensional characteristics usually disappear when setbacks and tapers are introduced. this is , when the pressure profile is different from the design and the drag is usually higher than expected, reducing the performance of the wing. Make it dark.

Figure 2 shows a wing placed at a sweepback angle A with respect to the direction perpendicular to the air velocity M- across the wing. A portion of the infinitely swept wing panel 11 is conceptually shown. According to the simple regression theory, the speed degree is decomposed into two components.

The first component is a component parallel to the wing panel, and can be expected to have an appreciable effect on the performance of the wing. I won't show it because I don't have it. The components of most interest are the leading edge 12 of the wing panel 11 (and the leading edge 1 2 is the vertical component of the air velocity M, perpendicular to the trailing edge 14). Only this MW determines the nature of the pressure field of the wing panel 11.

Therefore, the simple sweepback theory is based on the designer's belief that the direction from the wing in a straight line parallel to MN Analyze the airfoil shape, analyze the airfoil at a speed of M-cosA, determine the pressure region, and calculate the pressure Divide the force area by cjs''A to allow defining the pressure area of the wing panel 11 . This simple regression theory can work in two dimensions by the designer, but Unable to translate results to 3D environment via corners.

For example, extract the airfoil perpendicular to the recession line and obtain the two-dimensional state given by the simple regression theory. By analyzing and converting the resulting pressure back in the flow direction, the following equation holds: Ru.

Cps-o=Cpx-o cos”A (1) However, Cp is the pressure coefficient, Eight is the receding angle.

However, in the simple sweepback theory, the position of the shock wave in the wing panel 11, the sweepback angle Assume that is always the same and that it is connected to the geometric features. Therefore, simple The concept of sweepback theory is extended to finitely tapered wings where the sweepback angle of the wing varies between the leading and trailing edges. There must be.

FIG. 3 shows, in a finitely tapered wing 16, from a sweepback angle A1 at the leading edge 18 to a sweepback angle at the trailing edge 20. It conceptually shows the change in the receding angle up to angle A. The retreat of the shock wave directly above is the position of the wing. depends on. Airplanes now increase wing angle of attack for increasing Matsuha number When maneuvered, a shock wave develops and moves rearward on the wing, reducing the sweep angle and increasing the resistance increasing resistance.

Conventional design procedures typically determine the location of the spanwise generatrix 22 of the quarter chord line of the wing 16. It was called based on the assumption that determines the receding angle for two-dimensional/three-dimensional conversion. this geometry The scientific or flat sweep angle guidance does not take into account the aerodynamics of the wing, and the wing is When tested, develops deviations in performance from initial design criteria.

Conventional wing design and wing design according to the present invention have been developed from the initial stage of defining the performance requirements of the wing. It has started.

Once the performance requirements of the wing are optimally defined, conventional wing designs and the wing design of the present invention can be The next step in the plan is to size the wings.

Basic parameters of wing area, load, sweep angle, taper, wing thickness and define performance requirements. Aspect ratio is determined. After such a decision, the next step in conventional wing design is to determine the shape of a two-dimensional airfoil based on the defined parameters. analysis or details of the two-dimensional airfoil to achieve specific pressure characteristics by numerical analysis procedures. There are many guidelines for contour lines. These guidelines are known to those skilled in the art. It is. The usual discussion is A, B, Mr. Ines Notsuta's The Aeronxutic il Journal Volume 83 No. 819 March 1979'Computer- Aided Design: Aerodynamics”.

According to the first intention of the invention, the conditions under which the two-dimensional design is carried out are defined. be For the lift requirements of the wing, determine the lift requirements and design matsuka number for each airfoil shape. It is necessary. According to the invention, the airfoil is such that the effective retraction of the airfoil prevents the shock wave directly above the airfoil. The design is based on the discovery that location is constrained. This position allows the wing to ``Planned incorporation'' is considered the most advantageous for specific uses. This means that the wing is in use Eliminates the difficulty of conventional technology in designing an ideal airfoil shape without actually experiencing it. ing.

Therefore, according to the invention, the actual flight of the wing due to the actual desired position of the shock wave directly above An aerodynamic sweepback angle can be selected or defined.

After this aerodynamic recession angle is determined, the lunar velocity or Matsuha number in two dimensions is determined. is calculated by multiplying the three-dimensional velocity by the cosine of the aerodynamic sweep angle, as shown in the following equation: Ru.

M　2-D-M　S-OcosA　shschwav*　(2) Airfoil lift coefficient is calculated by dividing the three-dimensional lift coefficient by the square of the cosine of the aerodynamic sweep angle, as follows: I can't stand it.

CLz-o　-CL3-D/　eos”A　＊　ha　akwawe　(3) Subordinate Therefore, the two-dimensional airfoil is designed to provide a desired pressure range (shock wave position), as is known to those skilled in the art. Using trial and error, computer aided design (CAD) or iterative design procedures starting from It is designed with

For example, experience with the photographic design method shows that supercritical airfoils can be designed to meet the following rules of thumb: It is implied that it will be measured.

M,, +CL/10 + t/c -0, 95 (4) The standard answer is usually air power The mechanical influence coefficient determinant is expanded and inverted to obtain the boundary shape for the blade mean surface by is required to be multiplied by the column vector representing the state.

However, p = wing panel pressure difference, Aww - aerodynamic influence coefficient determinant, (dz/dx) c-airfoil/wing height inclination angle, α-angle of attack.

In the reverse mode, the program run may be distorted by 1°. Therefore, the influence coefficient The determinant is multiplied by the desired longitudinal and longitudinal load distributions before inversion. .

((dx/dx)c-σ)-[Aww]XPw (6) The required load is is the chordal loading of the airfoil scaled for sweep angle across the wing. The subsonic state is /<Nell method predictions were chosen to ensure that they are not unduly compromised by wave effects. Ru.

A second intention according to the invention is that the three-dimensional airfoil maintains the desired two-dimensional pressure area and thus How can a 3D wing be designed to produce the desired lift and minimum drag? It relates to being formed based on shape.

FIG. 4 shows how the two-dimensional airfoil is placed on the three-dimensional airfoil 16 so that the maneuver speed is determined by the airflow downstream. increases as the flow increases. Furthermore, the effective airflow component is obtained when the wing is cut in a straight line. It acts on the surface shape that is not visible. The equivalent two-dimensional airfoil shape for any wing is the number along the chord. This is generated by constructing a chord line perpendicular to the local recession line at each location. . The formula for configuring the effective blade shape uses the ratio of the curved chord C8-0 and the chord length in the airflow direction. You can get it.

as well as Y2-D-Y mtreamwis*(C!-D/Cst+*a+mw i we) (8) However, the horizontal axis of the point in the x-x-y plane, in the Y-X-Y plane The vertical axis of the point, C - Chord length of the wing cross section, Δ10. - the local sweepback angle of the spanwise generatrix, AB = sweepback angle at the leading edge, and ATt - Retraction angle at trailing edge.

Therefore, the three-dimensional shape of the wing 16 in FIG. 28B, 28C, and 28D of the wing 16. Arranging the airfoil shape of the wing along the curved line 24 that constitutes the intersecting oblique chord line. defined by

Typically, the required twist distribution is also required to be included in the wing profile. The swept wing is the root normally has a positive angle, but the wing tip has a negative angle with respect to the root.

After the airfoil shape has been defined by placing the airfoil on the airfoil, cutting in the airflow direction is within the skill of those skilled in the art. is carried out from the design shape in a known manner to determine the wing profile for the lift of the wing. death However, since the design airfoil is found along the curved line 24, the shock wave is located at a favorable location. As a result, the airfoil cross-sectional performance of a three-dimensional airfoil approaches that of a two-dimensional airfoil.

This means that the design airfoil has one specific direction generatrix, for example, a span of 1/4 chord length. Clear contours as opposed to traditional design procedures where wings are typically placed perpendicular to the generatrix. are doing.

It is understood that the design airfoil typically varies from the root of the wing to the tip of the wing. for example , the root airfoil has a thickness on the order of 12% and the tip airfoil has a thickness on the order of 8%. It may have a thickness. In this case, the above abbreviated design procedure may actually , so it is repeated at various locations along the wing span. Therefore, wings The shape must change along the wing span.1 For more complex wing body shapes, , it makes sense that the design procedure of the present invention is repeated at various points along the wing. be understood.

FIGS. 5A to 5F show correlations that make the design procedure according to the present invention effective as described above. Shows research. Effective two-dimensional airfoil 1, leading edge sweep angle of 40° and Equation 7 and Designed for an equivalent 2-D state with f2M = 0.689 and CL = 1.7 using Eq. The blade taper ratio shown in Fig. 5G is λ=C++p/Cr@@r. is extracted from the flat plate-shaped empirical wing 30. Table 1 shows that for chord 32 in the airflow direction. Sweeping angle and chord of the wing shock wave for each of the six cases of Figures 5A to 5F carried out along forming a position. These sweepback angles are the actual values used to convert in Eq. The aerodynamic sweep angle of is a two-dimensional Matsuha number. In each case, the airfoil lift coefficient normalized per sweep was required to match the lift coefficient of the wing. When there is no shock wave (M-0, 70) uses a 1/4 chord sweepback angle.

As shown in FIGS. 5A-5F, the pressure regime of an empirical three-dimensional airfoil and, for example, the present invention There is good agreement between the backward taper theory and the tied two-dimensional cone-Garabedian analysis. It's good.

According to a third aspect of the present invention, it is possible to improve the performance evaluation of designed three-dimensional airfoils. It is possible. Two-dimensional resistance using traditional design evaluation procedures including airflow simulation techniques has lift simply added to the other wing drag components to predict the overall wing drag. It is a fixed number. According to the invention, when the shock wave develops in a very severe condition, the wing It is recognized that moving the Ru. Therefore, the wing does not actually have a three-dimensional sweep angle, but the wing state is determined by the maneuvering speed. It has an aerodynamic sweep angle that varies with the strength of. The wing contour resistance CD is as follows. I can calculate it.

Cop-CCax-o Cats-o) cos3Ashaahwave (9) The overall resistance C0 is obtained from the following equation.

Co-Cos+Co++Cop (10) However, Cd! -D-two-dimensional drag coefficient, -C□, -D-two-dimensional frictional resistance coefficient, CD*=zero profile-lift drag, and C0, = lift induced resistance.

FIG. 7 shows the results achieved in the empirical wing by implementing the backward taper theory according to the present invention. This shows an improved resistance prediction. For a given lift coefficient, as discovered in wind tunnel experiments, The actual drag is what would be expected when simple sweep theory forms the airfoil profile drag level. Higher than. However, the swept-taper theory according to the present invention shapes the airfoil profile resistance level. Agreement between theory and experiment is valuable when

Even after the design procedure according to the invention is used, the calculated I; three-dimensional wing performance , the shock wave position (effective sweepback angle), even if it is no closer than expected performance than desired. More than one attempt may be required to find the optimal combination of pressure range and resistance. stomach. The design can be changed in the event of such an event. If Baur Garabedian ・Kohn-Jamson (BGKJ) code C or the above-mentioned paper by Hynes Analytical techniques such as If indicated, for example, the shock wave position or the design speed or the airfoil shape can be changed. follow the selected design before the actual prototype is built and tested in the wind tunnel. A number of iterative designs can be achieved.

The design procedure of the present invention ensures that the two-dimensional design performance is obtained in three-dimensional quality. May be used on a wide range of airplane wings.

Currently in service and planned for the future, excluding supersonic transport and supersonic range combat aircraft Almost all airplanes can operate at transonic speeds for most of their flight time, excluding takeoff and landing. Ru. Accordingly, the present invention provides wing designs for transport aircraft such as airliners and executive jets. , and extreme cases where the painter simply changes the design requirements in a known manner to operate in the transonic region. It may also be applied to fighter aircraft, which are easy to maneuver.

FIG, [ WE for the 2nd year's resolution prayer to Matsu Arikari FIG.7 ooOa: A experiment Number of pit workers, C0 international search report

Claims (7)

[Claims] 1. Define performance requirements for the wing and use this performance requirement to introduce normal wing dimensions death, determining an aerodynamic sweep angle of the wing with respect to the actual position of the shock wave on the wing; The two-dimensional Mach number for each airfoil cross section is the cosine value of the aerodynamic sweep angle in the three-dimensional Mach number. The two-dimensional lift coefficient is obtained by multiplying the three-dimensional lift coefficient by the aerodynamic sweep angle. The shape of the two-dimensional airfoil is calculated by dividing the cosine value of determined on the basis of the number and the two-dimensional lift coefficient requirement, The cross-sectional shape of the wing in three dimensions is determined by the local recession of the wing at a series of locations along the chord line of the wing. It is defined by arranging the airfoil shape of the wing along the curved structure of the oblique chord line that is perpendicular to the chord line. A method for designing a transonic flight wing characterized by: 2. Claims in which the aerodynamic sweepback angle is determined by the actual position of the shock wave. The method described in paragraph 1. 3. The wing calculates the twist for the wing as a function of distance from the root and this calculated Claim 1 comprising the step of changing the shape of the wing in three dimensions according to the friction. The method described in item 1 or 2. 4. The method may be performed at each of a plurality of positions from the root of the blade to the tip of the blade. The method according to any one of claims 1 to 3. 5. Evaluate blade performance in three dimensions using secondary flow simulation technology to improve predictability Claim 1: generating performance results and improving the wing according to the expected performance results. to the method described in any of paragraphs 4 to 4. 6. 6. The method of claim 5, wherein said evaluation and refinement steps are repeated. 7. The step of evaluating the performance is performed according to the following relationship: CDp=(Cd2-D-C df2-D) The resistance CD of the entire cos3Λshockwave is obtained from the following formula: Ru. CD=CDa+CDi+CDp However, Cd2-D=two-dimensional resistance coefficient, Cdf2-D=two-dimensional frictional resistance coefficient, CD0 = Zero contour - Lift resistance, and CDi = lift induced resistance, 7. The method according to claim 5 or 6, comprising the step of evaluating external resistance. 8. Before The step of arranging the airfoils in the airfoil is performed according to the following relationship: se/C2-D)=(X/C)[(1/cosΛlec(1+tanΛlec・ tanΛLE))]+(1-(X/C))[(1/cosΛlec(1+tan Λlec・tanΛTE))] and Y2-D=Ystreamwise (C2-D/Cstreamwise) However , X=horizontal axis of the point in the X-Y plane, Y=vertical axis of the point in the X-Y plane, C = Chord length of the blade cross section, Λlec = local sweepback angle of the spanwise generatrix, ΛLE = sweepback angle at the leading edge, and ΛTE = Sweeping angle at trailing edge. 8. A method according to any of claims 1 to 7, in which: