JP2009012686A - Supersonic type aircraft configuration for reduction of rear end sonic boom - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a supersonic type aircraft configuration capable of solving a problem of a trim held by a conventional cross section area distribution of George and Seebass and attaining a rear end low sonic boom even if a horizontal stabilizer of usual wing back arrangement is adopted. <P>SOLUTION: In the machine body shape of the supersonic type aircraft, equivalent cross section area distribution, i.e., the sum of machine body axial direction distribution of a cross section obtained by cutting the machine body shape by Mach plane determined by flying Mach number and a cross section area distribution obtained by converting machine axial direction distribution of lift to cross section area distribution equivalent to it has distribution having the maximum value on the midway without monotonously increasing it in the machine axial direction, and is provided with a portion where the equivalent cross section area is increased again without monotonously reducing it in the process that the equivalent cross section area is reduced from a position for taking the maximum value to a rear end of the machine body. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates to a fuselage shape that reduces the rear-end sonic boom of a supersonic aircraft.

  In general, a supersonic aircraft is required to suppress a sonic boom, which is an acoustic phenomenon affecting a structure such as a person, an animal, or a building on the ground at the time of supersonic flight in order to satisfy an environmental compatibility requirement. Sonic boom reduction methods have been studied for many years, and the most promising method is to change the shock wave generation pattern by devising the shape of the fuselage to reduce the sonic boom strength on the ground. The shock waves generated from each part of a normal supersonic aircraft are two strong shock waves, the nose and the tail, accompanied by the phenomenon that waves with large pressure fluctuations propagate faster in the atmosphere as they propagate through the atmosphere. It is integrated and observed as an N-type pressure wave with a large pressure rise of 2 degrees on the ground. The shock wave generated and propagated by the supersonic aircraft propagates in a cone shape and reaches the ground as shown in the lower left of FIG. The N-type waveform at that time is composed of a shock wave that is increased in pressure from the atmospheric pressure by the nose part at once, and a shock wave that is gradually reduced in pressure from the low pressure to return to atmospheric pressure by the tail part as shown in the upper left of FIG. By the way, Concorde's sonic boom, which is representative of supersonic passenger aircraft, is said to have a strength of 2 to 3 psf and a sound equivalent to a nearby lightning strike.

Since supersonic flight over land and air is restricted due to the noise problem caused by the sonic boom, it has become a subject of practical application of supersonic passenger aircraft. The sonic boom reduction method described above forms a low sonic boom pressure waveform that is not a normal N type by correcting the shape of the airframe to suppress the integration of shock waves. George and Seabass proposed a boom minimization based on the near-field theory, including the rear end wave, and the “minimum overpressure waveform” shown in the upper right of FIG. 9 and the “minimum shock waveform” shown in the lower right of FIG. Two types of pressure waveforms were studied, and Non-Patent Document 1 presented a theory that focused on the sum of the equivalent cross-sectional area distribution obtained from the cross-sectional area distribution and lift distribution of the aircraft forming the low sonic boom pressure waveform. Here, the cross-sectional area distribution of an aircraft means that the aircraft is cut by a Mach plane determined by the flight Mach number [a plane in which a normal vector is inclined upward with respect to the aircraft axis by an angle μ = sin ⁻¹ (1 / M)]. This is the distribution of the projected area of the measured cross-sectional area in the machine axis direction. The theory of George and Seabass notes that the first factor that causes the airframe to change pressure to the atmosphere is the shape of the fuselage, and the second factor is the reaction of lift received by the wings. Assuming that this is an equivalent cross-sectional area that covers all directions in the same way as the fuselage, the sum of the equivalent cross-sectional area distribution obtained from the cross-sectional area distribution and lift distribution of this aircraft is It is a theory that a low sonic boom can be realized when a predetermined distribution is shown. However, when the shapes having these equivalent cross-sectional area distributions are calculated by back calculation, the nose shape becomes blunt with a large body drag. After that, Darden presented a technique and a program in Non-Patent Document 2 to reduce the aircraft drag caused by this nose using the cross-sectional distribution of George and Seabass. In addition, in the patent document 1, the present inventor presented a nose shape that uses a non-axisymmetric fuselage to achieve both a low sonic boom and a low fuselage resistance.

The cross-sectional area distribution proposed by George and Seabass is a cross-sectional area distribution that monotonously increases from the nose to the rear end of the fuselage as shown in the second row from the top of FIG. On the other hand, since the cross-sectional area distribution of the aircraft becomes zero when there is no intersection between the airframe and the Mach plane in the vicinity of the rear end of the airframe, the cross-sectional area distribution at that position must be the equivalent cross-sectional area of lift. The equivalent cross-sectional area of lift at a certain aircraft axis position corresponds to the total lift from the nose to that position. If it is assumed that only the main wing bears the lift, the equivalent cross-sectional area distribution of the lift has a maximum value at the rear end of the main wing and a constant value behind it. Therefore, in order to realize the above-described monotonically increasing cross-sectional area, it is necessary to distribute the main wings to the rear end of the airframe as in the airframe example shown in the upper part of FIG.
However, if a large lift is generated at the rear of the aircraft, a nose-down moment is generated around the center of gravity of the aircraft, making it impossible to trim the aircraft (vertical balance). In order to trim such a low sonic boom fuselage, a horizontal tail (stabilizing plate) that is usually located behind the main wing is used in front of the main wing (canard), or As a tailless aircraft configuration, it is necessary to balance the lift distribution by distributing the main wing long in the axis direction. As an example of a low sonic boom aircraft adopting a canard configuration, Patent Document 2 adopts a canard and a reverse V-tail with the main wing taken to the rear end of the aircraft. Generally, a canard-type supersonic aircraft has a problem of unstable pitch-up (non-linear head-up) during low-speed take-off and landing. The plane shape of the main wing, which distributes the lift in the longitudinal direction, has a large wing area in view of structural strength, and must be a wing with a small aspect ratio, increasing the frictional drag and induced drag of the aircraft.
Japanese Patent Laying-Open No. 2005-178491 (Patent No. 3855064) “Method for Determining Body Shape of Supersonic Aircraft and Body Shape of Front Body” Published July 7, 2005 US Patent No. 6,824,092 Frankhn, WM, "Aircraft Tail Configuration for Sonic Boom Reduction." Published November 30, 2004 US Patent 3,647,160 Alperin, M. "Method and Apparatus for Reducing Sonic Booms." Issued March 7, 1972 Seebass, AR and George, AR, "Designand Operation of Aircraft to Minimize Their Sonic Boom," Journal of Aircraft, Vol. 11, No. 9, pp.509-517, 1974. Darden, CM, "Sonic-Boom Minimization With Nose-Bluntness Relaxation." NASATP-1348, 1979.

  The object of the present invention is to solve the problem of trim that the cross-sectional area distribution of George and Seabass has in the past, and to realize a supersonic aircraft shape that can realize a low sonic boom at the rear end even if a horizontal tail wing arranged at the rear of the main wing is adopted. It is to present.

The aircraft shape of the supersonic aircraft of the present invention was converted into a cross-sectional area distribution obtained by cutting the fuselage shape on the Mach plane determined by the flight Mach number, and a lift-axis distribution in the axial direction was converted to an equivalent cross-sectional area distribution. The equivalent cross-sectional area distribution, which is the sum of the cross-sectional area distributions, has a distribution that has a maximum value in the middle without monotonically increasing in the direction of the axis, and the equivalent cross-sectional area decreases from the position where the maximum value is reached to the rear end of the aircraft. In the process, a portion where the equivalent cross-sectional area increases again without monotonously decreasing is provided.
Specifically, it is corrugated by giving the back of the fuselage lower surface unevenness, by using a horizontal tail that gives the airfoil a reverse camber (downward convex cross-sectional shape), or by placing an engine nacelle at the rear of the fuselage. A near-field pressure waveform having one or more combinations of expansion-compression-expansion is formed at the rear, and the sonic boom is formed by setting the rear of the ground sonic boom pressure waveform as a multi-stage pressure rise instead of the normal N-type. The strength was reduced.
In addition, the rear surface of the fuselage has an uneven surface, or a horizontal tail that has a reverse camber (downward convex cross-sectional shape) on the airfoil, or a combination of multiple shapes that arrange the engine nacelle at the rear of the fuselage. In the rear part of the field pressure waveform, a waveform canceling part by a combination of expansion-compression-expansion is formed.

The supersonic aircraft fuselage of the present invention forms a near-field pressure waveform having a combination of expansion-compression-expansion one or more times at the rear of the waveform, thereby making the rear portion of the ground sonic boom pressure waveform a normal N-type. Instead of the multi-stage pressure rise, it solves the trim problem of the cross-sectional area distribution of George and Seabass, and realizes a low sonic boom at the rear end even if a horizontal tail is used behind the main wing. be able to.
The present invention eliminates the need for distributing lift to the rear end of the aircraft, unlike conventional low sonic boom aircraft, even with low sonic boom designed aircraft, so the conventional horizontal tail can be placed behind the aircraft Therefore, trimming the aircraft is easy. (Effects of improving flight characteristics)
Also, in the conventional low sonic boom design, the wing plane shape cannot be optimized to maintain the lift to the rear end of the fuselage and the drag increases, or the wing plane shape with poor trim characteristics and Although it has been unavoidable that the trim drag increases due to the main wing arrangement, it is possible to adopt the main wing plane shape with the optimum lift-drag ratio and the main wing arrangement with the optimum trim characteristics by using the present invention. (Aerodynamic characteristics improvement effect)
According to the present invention, control of the pressure wave generated at the rear of the fuselage is facilitated, and there is an effect that the robustness with respect to fluctuations in flight conditions and atmospheric conditions of the rear-end sonic boom characteristics can be improved as compared with the conventional low sonic boom design. (Robustness improvement effect of rear end sonic boom characteristics)

  In the conventional low sonic boom design method, the equivalent cross-sectional area distribution of the fuselage is uniquely determined by specifying the flight Mach number, flight altitude, fuselage weight, fuselage length and the target sonic boom pressure waveform. Lift distribution (main wing shape / arrangement) and cross-sectional area distribution must be determined to satisfy the distribution. Inevitably, the specified cross-sectional area distribution imposes strong constraints on the body volume of cabins, cockpits, fuel tanks, etc., and deprives design freedom. Conversely, to determine the low sonic boom equivalent cross-sectional area distribution so that the required volume can be secured, it is necessary to constrain mission requirements (upstream side of aircraft design) such as flight Mach number, flight altitude, and aircraft weight. It is difficult to design an efficient aircraft. Specifically, if the required volume exceeds the low sonic boom equivalent cross-sectional area distribution determined from the mission requirements, it is necessary to increase the equivalent cross-sectional area distribution by increasing the flight altitude, increasing the aircraft weight, etc. However, according to the present invention, it is possible to independently determine the mission requirements such as the flight Mach number, the flight altitude, the aircraft weight, and the required aircraft volume, which has the effect of increasing the degree of freedom in design. (Supersonic aircraft low sonic boom design freedom expansion effect)

In the present invention, in order to solve the problem of trim that the cross-sectional area distribution of George and Seabass has, a low sonic boom cross-sectional area distribution that is different from the cross-sectional area distribution proposed by George and Seabass is presented. By adopting the cross-sectional area distribution of the present invention, the main wing is arranged near the center of the airframe without being distributed to the rear end of the airframe, and the rear end low sonic boom property is achieved even if the horizontal tail wing arranged at the rear of the main wing is adopted. It becomes possible to hold.
The principle on which the present invention is based is shown in FIG. The pressure wave has a non-linear property that propagates faster in the air as the pressure rise is larger, and the positive pressure wave due to the combination of compression wave and expansion wave is forward, due to the combination of expansion wave and recompression wave Negative pressure waves shift backwards. Therefore, the pressure wave by the combination of compression wave + carving wave + recompression wave as shown in the upper part of FIG. 1 increases in length as it propagates through the air, and grows into an N-type wave while forming a shock wave. Go. Conversely, the pressure wave generated by the combination of expansion wave + compression wave + expansion wave as shown in the lower part of FIG. 1 cancels out the positive part of the waveform shifted forward and the negative part of the waveform shifted backward. One shock wave is formed and stays in place.

In the present invention, the above principle is applied, and the cross-sectional area distribution that takes the maximum value near the rear end of the main wing arranged near the center of the fuselage is shown by the solid line in the left diagram of FIG. As shown, the once reduced cross-sectional area is increased once again to give a bulge near the rear end of the aircraft (Method 1). In this cross-sectional area distribution, an expansion wave is generated when the cross-sectional area decreases from the position where the cross-sectional area is maximum, a compression wave is generated at the bulge near the rear end of the aircraft, and then an expansion wave is generated again. As a result, the pressure waveform as shown by the solid line in the center of FIG. The hatched portion in the figure is the canceling portion in the above principle. When the shock wave stays at this position, the rear-end sonic boom reaches the ground as a two-stage pressure rise without being integrated into the N-type wave. To do. (A solid line in the right diagram in FIG. 2) Examples of aircraft shapes for realizing such an equivalent cross-sectional area distribution are shown in FIGS. 3 (1), (2), and (3). In FIG. 3 (1), the combination of expansion wave-shock wave-expansion wave-shock wave is realized by providing irregularities on the lower surface of the rear body shape.
In FIG. 3 (2), a combination of shock wave-expansion wave-shock wave is realized by a horizontal tail instead of a bulge on the lower surface of the rear fuselage. At this time, by designing the shape of the aircraft to efficiently achieve a favorable sonic boom pressure waveform by adjusting the strength of the shock wave and expansion wave by increasing the blade thickness of the cross section of the horizontal tail or by attaching a reverse camber Is possible. FIG. 3 (3) shows an example of realizing a combination of shock wave-expansion wave-shock wave by arranging an engine nacelle instead of the horizontal tail.

  In the present invention, it is possible to divide the rear sonic boom pressure waveform into a plurality of small pressure increases by making the cross-sectional area distribution at the rear of the fuselage multi-stage unevenness (method 2). The solid line in the left figure of Fig. 4 shows an example in which the cross-sectional area distribution at the rear of the fuselage has three steps of unevenness (the shape of the aircraft that has a cross-sectional area increased twice when it decreases from the maximum cross-sectional area position). From an aircraft having such an equivalent cross-sectional area, as shown by the solid line in the central diagram of FIG. 4, a near-field pressure waveform composed of expansion wave-shock wave-expansion wave-shock wave-expansion wave-shock wave is formed. Since the portion indicated by the oblique line remains in place by cancellation, a sonic boom waveform having a three-stage pressure rise is formed at the rear of the waveform, as indicated by the solid line in the right figure of FIG. Examples of aircraft shapes for realizing such an equivalent cross-sectional area distribution are shown in FIGS. 5 (1), (2), and (3). In FIG. 5A, a combination of expansion wave-shock wave-expansion wave-shock wave-expansion wave-shock wave is realized by providing a plurality of irregularities on the lower surface of the rear body shape. In FIG. 5 (2), a combination of shock wave-expansion wave-shock wave is realized by a horizontal tail instead of the second bulge on the lower surface of the rear fuselage. At this time, by designing the shape of the aircraft to efficiently achieve a favorable sonic boom pressure waveform by adjusting the strength of the shock wave and expansion wave by increasing the blade thickness of the cross section of the horizontal tail or by attaching a reverse camber Is possible. FIG. 5 (3) shows an example in which a combination of shock wave-expansion wave-shock wave is realized by disposing an engine nacelle instead of the first-stage bulge.

  As the best mode for carrying out the present invention, FIG. 6 shows an example in which the present invention shown in FIG. 2 is applied to the rear of the equivalent cross-sectional area distribution of George and Seabass (the broken line in the left figure of FIG. 6). . The equivalent cross-sectional area distribution indicated by the solid line in the left diagram of FIG. 6 is an equivalent cross-sectional area distribution determined by the optimization design tool so as to minimize the negative pressure peak at the rear of the sonic boom pressure waveform. The distribution to be identified. The near-field pressure waveform estimated by applying the linear panel method to the axisymmetric object having the equivalent cross-sectional area distribution is shown by a solid line in the center diagram of FIG. Similar to the near-field pressure waveform shown by the solid line in the center of FIG. 2, a positive pressure peak is formed at the rear of the waveform due to the shock wave from the convex portion at the rear of the object. The ground sonic boom pressure waveform estimated by the modified linear theory from the near-field pressure waveform of the solid line in the center of FIG. 6 is shown by the solid line of the right-hand side of FIG. It can be seen that the rear portion of the sonic boom waveform is formed by a two-stage pressure increase.

  FIG. 7 shows a preferred embodiment of the present invention shown in FIG. The equivalent cross-sectional area distribution indicated by the solid line in the left diagram of FIG. 7 is the equivalent cross-sectional area distribution determined by the optimization design tool so as to minimize the negative pressure peak at the rear of the sonic boom pressure waveform. The distribution is within the scope of the claims. The near-field pressure waveform estimated by applying the linear panel method to the axisymmetric object having the equivalent cross-sectional area distribution is shown by a solid line in the center diagram of FIG. Similar to the near-field pressure waveform shown by the solid line in the center of FIG. 4, a plurality of positive pressure peaks are formed at the rear of the waveform due to the shock wave from the convex portion at the rear of the object. The ground sonic boom pressure waveform estimated by the modified linear theory from the solid line near-field pressure waveform in the center of FIG. 7 is shown by the solid line on the right side of FIG. It can be seen that the sonic boom waveform rear part is formed by multi-stage pressure rises by canceling each other.

  In FIG. 8, the example which implemented this invention in the design of a three-dimensional airframe shape is shown. As shown in Fig. 8 (a), the main wing is located at the fuselage center and the lift distribution cannot be maintained up to the rear end of the fuselage. As shown by the solid line in the left figure of FIG. 8 (b), it is possible to form a re-increasing part (convex part) of the equivalent cross-sectional area distribution at a place where the equivalent cross-sectional area decreases at the rear part of the airframe. The near-field pressure waveform estimated by applying the linear panel method to the three-dimensional airframe shape is shown by a solid line in the center diagram of FIG. It can be seen that a positive pressure peak is formed at the rear of the waveform due to the shock wave from the convex portion at the rear of the object. The broken line in the figure shows the near field calculated by applying the linear panel method to the axisymmetric object having the equivalent cross-sectional distribution of George and Seabass shown by the broken line in the left figure of FIG. 8 (b) for comparison. It is a pressure waveform. The ground sonic boom pressure waveform estimated by the modified linear theory from these near-field pressure waveforms is shown in the right figure of FIG. 8 (b). When the portions shown by the oblique lines in the center figure of FIG. 8 (b) cancel each other, It can be seen that the rear portion of the sonic boom pressure waveform indicated by the solid line is formed by a two-stage pressure increase.

  The present invention is used in the design of a supersonic aircraft, when the main wing is placed in the center of the fuselage and the lift distribution cannot be maintained (or not desired) up to the rear end of the fuselage, and the rear sonic boom needs to be reduced. it can.

It is a figure explaining the principle of this invention. It is a figure explaining the technique 1 of this invention. It is a figure explaining the form which implement | achieves the method 1 of this invention. It is a figure explaining the method 2 of this invention. It is a figure explaining the form which implement | achieves the method 2 of this invention. It is a figure explaining 1st Example of this invention. It is a figure explaining 2nd Example of this invention. It is a figure explaining 3rd Example of this invention. It is a figure explaining the generation | occurrence | production principle of a N-type waveform, and the conventional low sonic boom. It is a figure explaining the low boom design method of Darden.

Claims (7)

  Equivalent cross-sectional area, which is the sum of the cross-sectional area distribution of the cross-sectional area cut by the Mach plane determined by the flight Mach number and the cross-sectional area distribution obtained by converting the axial distribution of lift to the equivalent cross-sectional area distribution In the process where the distribution has a maximum value in the middle without monotonically increasing in the direction of the axis and the equivalent cross-sectional area decreases from the position where the maximum value is reached to the rear end of the aircraft, Aircraft shape of supersonic aircraft, characterized in that it has an area that increases in area.   Sonic is formed by forming a near-field pressure waveform having a combination of expansion-compression-expansion one or more times at the rear of the waveform, and by setting the rear of the ground sonic boom pressure waveform as a multi-stage pressure increase instead of the normal N type The body shape of the supersonic aircraft according to claim 1, wherein the boom strength is reduced.   The body shape of a supersonic aircraft according to claim 2, wherein a corrugation canceling portion by a combination of expansion-compression-expansion is formed at the rear portion of the near-field pressure waveform by providing irregularities on the lower surface of the rear portion of the fuselage.   The airframe shape of a supersonic aircraft according to claim 2, wherein a wave canceling portion by a combination of expansion-compression-expansion is formed at the rear of the near-field pressure waveform by the horizontal tail.   The horizontal tail wing shape has a reverse camber (downward convex cross-sectional shape) to generate a strong shock wave-expansion wave-shock wave below the fuselage, and a waveform by a combination of expansion-compression-expansion at the rear of the near-field pressure waveform The body shape of the supersonic aircraft according to claim 4, wherein a canceling portion is formed.   The aircraft shape of a supersonic aircraft according to claim 2, wherein an engine nacelle is disposed at the rear of the fuselage to form a waveform canceling portion by a combination of expansion-compression-expansion at the rear of the near-field pressure waveform.   A body shape of a supersonic aircraft that forms a waveform canceling portion by a combination of expansion-compression-expansion at a rear portion of a near-field pressure waveform by combining a plurality of shapes described in claims 3 to 6.

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