RU2351503C2 - Integral configuration aircraft - Google Patents

Abstract

FIELD: aircraft engineering.

SUBSTANCE: invention relates to aircraft engineering. Proposed aircraft consists of empennage, control surfaces, power plant, aircraft systems, equipment, fuselage, wing with the sweepback of leading edge of 30° to 35° and leading-edge wing extension sweepback of 60° to 70°. The section of transition from wing to fuselage is formed by a symmetric section with the thickness of 13% to 14%. The outer wings are formed by unified supercritical section with thickness making from 13% to 10% with linear section thickness variation over the wing spin. The section twist on inboard rib is negative approximating to -3°, that of the transition from leading-edge wing extension to wing spin is positive approximating to +2°. Note that the law of twist angle variation between the above sections is taken to be linear.

EFFECT: higher stability and safety.

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Description

The invention relates to aircraft and can be used to design integrated structures of aircraft of various types and purposes.

There are known schemes of integrated aircraft with a single fuselage and wing structure, which accepts the main loads with all the elements (see the encyclopedia Aviation (edited by GP Svishchev, Russian Encyclopedia Publishing House, Moscow, 1988). In practice, the integral the design is used on aircraft made according to the “flying wing” scheme, or for multi-mode combat aircraft: domestic Tu-160 or American B-1, which are selected as a prototype of the proposed solution (V.G. Rigmant. “Aircraft Design Bureau A.N. Tupolev. "- M.: Rusavia, 2001 ., Str.244-249).

The main structural feature of such an aircraft is that the spars and chords located in the conditional root section have a large height, which requires a correspondingly longer length, which makes it necessary to use large influxes of very large sweep in the layout, with which you can go to the base (trapezoidal ) wing.

With the accepted and well-known methods of forming the wing system, it is required to provide direct twist of the wing sections from positive (about + 2 °) swirl angles of the root profile to negative (about -2 °) swirl angles of the end profile, usually with a linear change in the swirl angles between these sections. Thus, the problem of maintaining a stable flow flow on the wing to the maximum achievable angles of attack is solved, and tear-off phenomena begin at the side of the fuselage in a limited area, since other profiles along the wing span have not yet reached critical angles.

However, the large size of a single root chord of an integrated aircraft does not allow to fully take advantage of the generally accepted laws of changing the angles of twist of the sections. With the appearance of a gap in the influx, the longitudinal moment at the pitching sharply increases, the aircraft becomes unstable and unsafe. It is required to find another law of twisting of airplane profiles, which, on the one hand, preserves as much as possible the general idea of the separation propagation in the direction from the root to the end section, on the other hand, eliminates the onset of tearing phenomena on the influx and the appearance of large longitudinal moments for cabling in the operating range angles of attack.

To achieve this goal, a unified wing-fuselage system is proposed, in which the consoles sweep along the leading edge is 30-35 and the influx sweep is 60-70 °, a single cross section is formed by a symmetrical profile with a thickness of 13-14%, the consoles are formed by a single supercritical profile with a thickness of 13 to 10%, with a linear change in profile thickness over the range, while the curl of the profiles on the side rib is negative (up to -3 °), on the transition line of the influx-console is positive (up to + 2 °), on the side rib is negative (up to -3 °), and zako the twist angle changes between the specified sections of the adopted line.

The essence of the proposal is illustrated by drawings.

Figure 1 shows a diagram of an integrated aircraft. Figure 2 - the results of purging the model with the generally accepted pattern of swirling profiles on the wing. Figure 3 is a new law of changing the curl of the profiles in a single system. Figure 4 - experimental data obtained by blowing a new model in a wind tunnel.

The plane 1 (Fig. 1) consists of an integral unit 2, which includes the front part with the crew cabin 3, the fuselage 4, which passes into the wing 5, which consists of an influx 6 and the console 7. The plane has a vertical 8 and horizontal 9 tail unit with rudders 10 and 11. The power plant consists of engines 12. Aircraft and engine systems, equipment and devices that are not shown conditionally are installed on the aircraft.

All flight modes and stages on an integrated circuit aircraft are performed without restrictions and do not require special skill.

But at large angles of attack due to flow separation at the beginning of the influx, a significant longitudinal moment appears (Fig. 2), which increases with increasing angle of attack and becomes unacceptable.

Experimental and theoretical studies have shown that the reason for this phenomenon is that, using the generally accepted scheme for changing the torsion angle of the profiles along the wing span (Fig. 3, line “A”), which really helps to improve the behavior of the console at large angles of attack (stall comes from the root section and goes slowly along the span), we get an unacceptable longitudinal moment due to the large size of the root chord, even with a slight stall. In the integrated circuit, it is necessary to change the position of the beginning of the breakdown and move it to a position where the length of the chord is small. Moreover, it is advisable to use a supercritical profile with moment characteristics in these wing sections, which will balance the longitudinal longitudinal moment in these sections. It was determined that in the best way this problem can be solved by applying the law of variation of the swirl angles in scope, which is shown in figure 3 (line "B"). At the same time, negative swirling of the profiles on the side rib (up to -3 °), positive swirling on the influx-cantilever transition line (up to + 2 °) and negative swirling on the side rib (up to -3 °) are ensured. The law of variation of the torsion angles between the indicated base sections is adopted linear.

Figure 4 shows the new characteristics of an integrated aircraft when applying the new law of twist profiles. Obviously, in the entire investigated range of angles of attack, the aircraft is stable and has high aerodynamic characteristics.

Claims (1)

Aircraft integrated circuit, consisting of plumage, rudders, power plant, aircraft systems, equipment and a single fuselage design with a wing with a sweep along the leading edge of 30-35 ° and an influx with a sweep of 60-70 °, containing supercritical and symmetric profiles, characterized in that a single cross section is formed by a symmetric profile with a thickness of 13-14%, the arms are formed by a single supercritical profile with a thickness of 13 to 10% with a linear change in profile thickness over the span, while the curl of the profiles on a single b the rib rib is negative (up to -3 °), on the influx-cantilever transition line is positive (up to + 2 °), on the side rib is negative (up to -3 °), and the law of changing the torsion angles between the indicated sections is accepted linear.

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