Aircraft and its swept wing

FIELD: transport.

SUBSTANCE: set of invention relates to aircraft engineering. Aircraft comprises airframe, swept wing, tail unit and jet engine. Airframe is characterised by selection of coordinates of outer surface outline points. Swept wing comprises cantilever parts and wing center section defined by coordinates of the upper and lower outlines of aerodynamic profiles located in wing basic sections.

EFFECT: reduced weight, higher comfort for passengers.

6 cl, 27 dwg, 13 tbl

 

The claimed group of inventions relates to an aircraft, namely the technical solutions of aircraft intended primarily for in-service regional airline, and swept wings for this type of aircraft.

Known technical solutions planes (see, for example, patents for utility models of the Russian Federation No. 13647 and 19814), which include the fuselage, wings, vertical and horizontal tail. The fuselage of the aircraft in accordance with these technical solutions are composed of sequentially placed nose and tail compartments. Wing in these decisions connected with the nose compartment, and the vertical and horizontal tail with the tail compartment. The technical problem to be solved with these useful models is improving aerodynamic performance of the aircraft while maintaining optimal dimensions for placement in the forward compartment 2...4 seats for crew and passengers. This technical problem is solved in these useful models by optimal choice of the geometry of the outer surface of the skin of the fuselage of the aircraft, which is defined by the coordinate sets of its points. Point on the surface of the fuselage selected from the contours formed by sections of the fuselage transverse planes, each path defined by three points: the first and second points taken as the point of intersection of the contour with the PLO is the bone of symmetry of the aircraft, and as a third point taken contour point, most remote from the plane of symmetry of the aircraft.

The technical problem in these utility model is solved for very small light aircraft, such as aircraft for agricultural purposes or aircraft for local airlines. These solutions cannot be used for the development of regional aircraft higher dimensions and to develop a model number of regional aircraft of different Seating capacity.

Known technical solution of a passenger plane Tu-134 (see, for example, Vairocana, Tu-134, ed. "Transport", M, 1972), the diameter of the cross section of which is of 2.9 m and the length of 37.1 m, while the lengthening of the fuselage (the ratio of the length of the fuselage diameter) is λ=12,8. The configuration of the cross-section of the fuselage allows you to put in the passenger cabin 76 passengers, while on each side of the plane are two rows of chairs. Used in this technical solution, the cross-section of the fuselage does not allow development of a model range of passenger aircraft, which includes various passenger aircraft: aircraft, designed to carry 76 passengers, aircraft, designed to carry, for example, 60 and 95 passengers with a high degree of unification, cover the soup using a single for all aircraft solutions of the cross-section of the fuselage and wing of the same geometry, leads to deterioration of the aerodynamic characteristics of the aircraft and significant weight cost. In addition, due to small (less than 3 m) transverse diameter of the fuselage level of comfort of passengers does not meet modern requirements. Do not meet modern requirements and some of the characteristics of the cargo compartment.

Known technical solutions swept wings haul aircraft with cruising speeds of 800-850 km/h, for example, the technical solution of the wing passenger aircraft Airbus industry a-320 (see Passenger aircraft Airbus industry a-320, ed. Zaitseva, N.N., p.21-23, TSAGI, 1993), the technical solution of the swept wing, described in the patents of the Russian Federation 1827975 (EN 1827975, VS 3/00, 1995) and 2228282 (EN 2228282, VS 3/14, 2002).

So the swept wing in accordance with patent RF 1827975 made with the elongation λ>7, sweep χ°=20-35° and has moving parts, starting with 20% of poluraspada profiles with relative small thickness=7...13% and the median lines of positive curvature with respect to the maximum curvature to its average value f/fcp=1,4...1,5, with the median line profiles of the moving part of the wing is designed so that the product of the square of maximum curvature on the average value of the curvature is the ratio of Xf2*fcp=4000...5000, and the angle between the y chord section and the tangent to the median line at the end point of the profile has a value of less than 6°.

The swept wing in accordance with patent RF 2228282 includes aerodynamic bearing surface formed on the basis of non-planar middle surface, smoothly varying along the span of the wing. During the transition from the side section to the wing tip S-shaped median surface with negative concavity in the caudal parts of the aerodynamic profiles gradually disappears, and its border with areas of positive concavity gradually shifted in the caudal part of the aerodynamic profile.

However, these technical solutions have the following disadvantages.

First, these solutions do not take into account the peculiarities of the wing flow associated with the influence of consoles installed under wing engines, nacelles and pylons. Secondly, significant technological difficulties in the manufacture of swept wings in production, as often the manufacturer covering the wing is fully consistent with the complex laws of change of the median surface and other characteristics of the aerodynamic surface of the wing, leads to the necessity of making a covering of the wing in the form of a surface of double curvature.

The closest analogue to the proposed technical solution "Airplane" is a decision on the patent of the Russian Federation No. 2277058 (publ. 27.05.2006,), which solves the problem of the development of sa is Oleta, designed for mainline transportation services, with variable capacity. In accordance with this decision, the aircraft includes a fuselage, United with him swept wing, vertical and horizontal tail. The fuselage is made up of sequentially placed nasal compartment (in the terminology of the source - front compartment, a front compartment (in the terminology of the source - removable front cover), the Central compartment, the rear compartment (in the terminology of the source - removable rear compartment) and the tail compartment. In this solution, the swept wing is connected with the Central compartment and comprises a center section and outer wing panels. In addition, this technical solution swept wing includes two transitional compartment, which are, if necessary, between the center section and outer wing consoles. Vertical and horizontal tail fixed in this solution in the tail section. Under consoles wing on pylons mounted jet engines.

In this solution, the nose and tail compartments made in the form of combinations tapering to the nose or tip of the fuselage forms a cylindrical form. Front and Central compartments and the back and tail compartments are joined to each other by a cylindrical forms, with the plane of the junction of the nose and the front and rear compartments and tail compartments held the Yat on the cylindrical part of the fuselage.

Change in passenger aircraft in this technical solution is achieved by changing the length of the fuselage by removing or installing the front and rear compartments, in combination with the change of the area of the swept wing by appropriate selection of the mentioned transition compartments of the wing.

Using this technical solution is justified for long-haul passenger aircraft, designed to carry more than 200 passengers. However, when using this technical solution for the development of a model series of aircraft of smaller dimension, designed to provide regional traffic, technically and economically not justified, because the change in passenger aircraft connected not only with a significant readjustment of production due to the changing geometry of the wing, but with the large amount of aerodynamic, structural and flight test with a modified geometry of the swept wing.

The closest analogue to the proposed technical solution of the swept wing is the solution of the wing (see engineering, encyclopedia, volume IV-21, book 1, Aerodynamics, flight Dynamics and strength, p.84-85, RIS, ed. "Engineering", M., 2002), which contains a console and centreplane part. Each of the outer wings are formed by sections of surfaces of single and double is rivini, smoothly mated with five aerodynamic profiles. The shape of each profile is composed of upper and lower procontrol. Airfoils are placed in the base sections, parallel to the plane of symmetry of the aircraft, and the plane of the first profile is placed in the on-Board cross-section. The front edges of each subsequent profile is shifted against the direction of flight relative to the leading edge of the previous profile. This technical solution at the transition from the on-Board cross-section to the wing tip decreases the relative thickness of the aerofoil. In addition, this technical solution at the transition from the on-Board cross-section to the wing tip decrease the angles of geometric twist of the wing sections.

The above-described analogs of the swept wing, including the nearest equivalent, are not intended for use in the development of a range of aircraft, including several different aircraft passenger capacity with the same size in cross section of the fuselage and its length, which is determined by the capacity of the particular aircraft. In addition, a typical solution scheme the aerodynamic design of the swept wing closest analogue does not take into account the need for accommodation in the immediate vicinity of the outer wings of the engine nacelles and pylons DWI is atela, that is typical for regional aircraft.

The technical problem solved by the claimed solution "Airplane"is the development of the aircraft, allowing for the creation of the model series of aircraft designed primarily for the maintenance of regional air services, by changing its Seating capacity without changing the geometric dimensions of the wing, in combination with high aerodynamic characteristics of each aircraft model range, reducing the weight of the aircraft structure and ensure a comfortable accommodation of passengers.

Problem solved claimed swept wings, is the development of a single wing geometry, providing high aerodynamic characteristics when used in a range of different aircraft passenger capacity in combination with the possibility of placing under the wing of two engines with large dimensions in the range of cruising flight modes, characterized by the number M=0,75...0,82. Problem solved claimed swept wings, is also developing aerodynamic wing, allowing to simplify the manufacturing technology detachable wing consoles.

The goal of the project claimed the decision "Airplane" is solved as follows.

Known solution of the plane containing the fuselage, United with him page is lavigne wing, vertical and horizontal tail. In addition, the aircraft contains jet engines mounted on pylons under the consoles. The fuselage of the aircraft in a known solution composed of sequentially placed bow, front, center, rear and tail compartments.

In the inventive solution, the new aircraft is that the outer surface of the fuselage specified coordinates in a rectangular coordinate system, the x-axis which is combined with the construction of the horizontal fuselage of the aircraft, the y-axis located in the plane of symmetry of the plane and the axis of applicat perpendicular to the plane of symmetry of the aircraft. Point on the surface of the fuselage selected from the contours formed by sections of the fuselage transverse planes. Each contour is specified by seven points, the first and second points taken as the point of intersection of the path with the plane of symmetry of the aircraft, as a third point taken contour point, most remote from the plane of symmetry of the aircraft. As the fourth and fifth points taken points on the path, away from the plane of symmetry of the aircraft by half, and the sixth and seventh points on seven-eighths of the distance between the third point and the plane of symmetry of the aircraft.

The surface of the nose compartment of the fuselage in the present decision are specified in the coordinate system, the x-axis is toroi directed against flight the origin is placed in the transverse plane containing the tip of the nose of the plane, with coordinates of points on the surface of the nasal compartment defined by the parameters given in table 1.

Table 1
X11Y111Y112Y113Z113Y114Y115Y116Y117
0-980,0-980,0-980,00,0-980,0-980,0-980,0-980,0
15-885,0-1061,7-980,0107,6-895,9-1050.8-930,2-1019,6
50-01,2 -1127,6-979,7197,9-820,7-1107,8-885,0-1051,3
100-719,4-1186,5-978,7281,9-747,0-1158,7-839,7-1079,3
300-495,4-1327,2-968,5496,7-540,6-1279,1-703,9-1142,1
600-248.5-1454,4-936,9714,4-305,1-1385,1-530,0-1187,4
1200150.3-1613,0-827,81024,7to 85.2-1507,8-200,6 -1207,9
1800650,4-1714,8-679.11242,0508,3-1576,0108,3-1180,5
24001081,9-1785,0-513,01400,0914,9-1614,6spreads for about 319.2-1128,8
30001338,2-1833,5-349,31516,11167.3-1634,6509,5-1067,8
34951481,7-1860,8-228,51588,11297,6-1642,1644,8-1018,7

The portion of the surface of the anterior compartment of the fuselage adjacent to the junction of the front compartment with the nose compartment, specified in the coordinate system, the x-axis which is directed against the flight, the origin is placed in the transverse plane, combined with the plane of junction of the nose and front compartments, while the coordinates of the surface points of the specified portion of the nose compartment defined by the parameters given in table 2.

Table 2
X21Y211Y212Y213Z213Y214Y215Y216Y217
01481,7-1860,8-228,51588,11297,6-1642,1644,8-1018,7
6001601,5-1881,2-112,11652,71401,1-1644,8742,1-969,4
12001677,3-1889,6-37,51697,01462,1-1646,4792,9-940,2
18001718,2-1890,0-4,91722,51491,4-1646,7821,6-927,0
24001730,0-1890,00,01730,01498,2-1647,5837,5-924,4

Part of the front surface of the compartment adjacent to the Central compartment, specified in the coordinate system, the x-axis which points in the direction of flight, origin combined with the plane of junction of the front compartment to the Central compartment, while the coordinates of the points of the surface of this part of the front compartment defined by the parameters given in table 3.

Table 3
X22/sub> Y221Y222Y223Z223Y224Y225Y226Y227
01730,0-2069,90,01730,01498,2-2062,5837,5-1851,9
6001730,0-2036,70,01730,01498,2-2028,2837,5-1746,3
12001730,0-1998,10,01730,01498,2-1982,3837,5-1460,8
18001730,0-1957,7 0,01730,01498,2-1901,1837,5-924,4
24001730,0-1917,80,01730,01498,2-1762,7837,5-924,4
30001730,0-1890,00,01730,01498,2-1647,5837,5-924,4

The surface of the Central compartment of the fuselage is specified in the coordinate system, the x-axis which is directed against the flight, and the origin of coordinates aligned with the plane of junction of the Central compartment to the front compartment, while the coordinates of the points of the surface of the Central compartment defined by the parameters given in table 4.

Table 4
X11Y311Y312Y313 Z313Y314Y315Y316Y317
01730,0-2069,90,01730,01498,2-2062,5837,5-1851,9
6001730,0-2094,50,01730,01498,2-2087,4837,5-1903,7
12001730,0-2109,20,01730,01498,2-2102,2837,5-1932,7
18001730,0-2115,20,01730,01498,2-2108,28375 -1946,0
24001730,0-2113,70,01730,01498,2-2106,8837,5-1944,8
30001730,0-2103,60,01730,01498,2-2096,6837,5-1935,1
36001730,0-2082,70,01730,01498,2-2075,6837,5-1913,8
42001730,0-2052,30,01730,01498,2-2044,9837,5-1878,6
48001730,0-2017,10,01730,0 1498,2-2008,3837,5-1824,4
54001730,0-1981,70,01730,01498,2-1966,6837,5-1738,7
55001730,0-1975,90,01730,01498,2-1958,6837,5-1718,4

The portion of the surface of the rear compartment, adjacent to the Central compartment, specified in the coordinate system, the x-axis which is directed against the direction of flight, and its origin is aligned with the plane of junction of the rear compartment to the Central compartment, while the coordinates of the surface points of this part of the rear compartment defined by the parameters given in table 5.

Table 5
X41Y411Y412Y413 Y413Y414Y415Y416Y417
01730,0-1975,90,01730,01498,2-1958,6837,5-1718,4
6001730,0-1942,60,01730,01498,2-1894,0837,5-1522,8
12001730,0-1910,60,01730,01498,2-1797,9837,5-924,4
18001730,0-1890,00,01730,01498,2-1679,7837,5-924,4
19021730,0-1890,00,01730,01498,2-1647,5837,5-924,4

The portion of the surface of the rear compartment of the fuselage, adjacent to the junction of the rear compartment with the tail compartment, specified in the coordinate system, the x-axis which points in the direction of flight, and the origin of coordinates is placed in the transverse plane, combined with the plane of junction of the rear compartment with the tail compartment, while the coordinates of the surface points of this part of the rear compartment defined by the parameters given in table 6.

Table 6
X42Y421Y422Y423Z423Y424Y425Y426Y427
01699,1-1304,9 197,11502,01497,9-1103,6924,2-530,0
6001721,0-1446,4120,21600,71506,5-1232,7895,2-622,4
12001729,5-1574,656,61672,91505,3-1348,8866,5-701,8
18001730,0-1686,313,61716,31500,0-1452,1844,6-775,6
24001730,0-1777,70,11729,91498,2-1538,7837,5-839,5
3000 1730,0-1844,60,01730,01498,2-1603,2837,5-888,6
36001730,0-1882,60,01730,01498,2-1640,2837,5-918,0
39971730,0-18900,01730,01498,2-1647,5837,5-924,4

The surface of the caudal compartment of the fuselage is specified in the coordinate system, the x-axis which is directed against the flight, and the origin of coordinates aligned with the plane of junction of the tail and rear compartments, while the coordinates of the points of the tail compartment defined by the parameters given in table 7.

-670,8
Table 7
X51Y511Y1 2Y513Z513Y514Y515Y516Y517
01699,1-1304,9197,11502,01497,9-1103,6924,2-530,0
6001665,2-1152,6350,51381,81483,1-963,5976,8-424,3
12001621,4-991,8606,31247,81463,5-819,81062,2-329,5
18001568,9-823,9845,81110,81435,61140,7-234,5
24001507,4-650,3949,5981,11399,2-514,61169,9-127,8
30001439,2-473,8945,4859,01355,9-353,51157,5-10,7
36001370,8-297,2940,3736,91306,6-192,31138,0106,6
42001302,4-120,6935,1614,81252,1-31,11111,0223,9
48001233,856,1928,5 491,91193,0130,21076,4341,4
54001125,9253,8869,2341,21091,8309,6994,3467,6
5874,5909,9523,2734,5175,3886,8548,9820,9622,6

In the inventive solution of the abscissa and ordinates of the points of each contour, the outer surface of the skin of the fuselage sections and their parts are calculated based on the following General proportions:

xij=Xij±Δ,

yijk=Yijk±Δ,

and applicati third points of the contours ratio

zij3=Zij3±Δz.

In the relations Xij, Yijk, Zij3- parameters of the coordinates of the outer surface of the fuselage skin above in tables 1-7, i - the index of the compartment, with i=1 for the nose compartment, i=2 for the front, where i is the index of the compartment, with i=1 DL the nasal compartment, i=2 for the front compartment, i=3 for the Central compartment, i=4 for the rear compartment, i=5 for the tail compartment, j=1 for the nose compartment, front compartment, adjacent to the nose compartment, the Central compartment of the rear compartment, adjacent to the Central compartment, and the tail compartment, j=2 for the front part of the compartment adjacent to the Central compartment, and the rear part of the compartment adjacent to the tail compartment, k is the number of points of the contour (k=1...7), and Δ, Δ and ∆ - values, values which do not exceed 5 mm

The surface of the cantilever part referred to the swept wing of the proposed solutions are made of sequentially placed along poluraspada wing, starting from the side section is smoothly conjugate to each other in the first area of the surface of single curvature, the first portion of the surface of double curvature, the second portion of the surface of single curvature and the second surface of double curvature. In addition, the surface of the cantilever part of the wing smoothly paired with aerodynamic profiles, one of which is placed in the on-Board cross-section of the fuselage, offset from the plane of symmetry of the aircraft at 1640±Δ7 mm, and its leading edge is offset from the junction of the Central and anterior compartments of the fuselage forward flight at 480±Δ mm, where ΔZ and ΔY values, values which do not exceed 100 and 30 mm, respectively.

In the inventive resh the research institutes mentioned jet engines placed under the first sections of double curvature consoles swept wing.

In addition, in the inventive solution the horizontal and vertical tail can be performed on most of them in the form of plots of single curvature, the surface of which is formed by two boundary symmetrical aerodynamic profile and line segments connecting the contours of the edge of an aerofoil, and the ends of the segments are placed in the y-paths of the same name percent of the chord of an aerofoil. The first plane of the airfoil vertical plane may be shifted from the construction of the horizontal fuselage at 1540±ΔY mm, and the second on 6810±ΔY mm, the plane of the first airfoil horizontal tail may be removed at 1100±ΔZ mm from the plane of symmetry of the aircraft, and the second on 5170±ΔZ mm, and the point of the front edge of the first airfoil horizontal tail can be shifted up to 750±ΔY mm from the construction of the horizontal fuselage. The horizontal stabilizer can be performed with an angle transverse V on the front edge of 3.5°...4,5° and sweep 34...35 degrees, and the vertical tail with the sweep of the leading edge 39...40 degrees. In the inventive solution, the contours of the first and second aerodynamic profiles of the vertical and horizontal tail can be set by parameters given in table 8.

Table 8
X0Y0
0,00,0
0,250,697
0,50,987
11,373
32,236
52,770
7,53,256
103,631
154,191
254,823
304,957
354,991
404,931
504,530
603,844
702,972
801,998
900,992
1000,0

The abscissa and ordinate of the contours of the aerodynamic profiles of the vertical and horizontal tail when this is calculated on the ratios

x=X0*/100,

y=Y0*/100±ΔP,

where X0, Y0- parameters of the aerodynamic contours of the profiles listed in the following table 8, the length of the corresponding chord of the airfoil, equal to the first and second aerodynamic profiles of the vertical fin 4496±Δ mm and 1364±Δ mm, respectively, and first and second aerodynamic profiles of the horizontal tail 2548±Δ mm and 815±Δ mm, respectively. In the preceding ratios, ΔY, ΔZ, ΔP, Δ - values, values which do not exceed 30, 30, 5 and 40 mm, respectively.

In addition, in the inventive solution of the plane sock first aerodynamic profile vertical plane can be offset from the plane of the junction of the tail and the rear cover forward in the direction of flight 951±Δ mm, and the toe of the first airfoil horizontal tail - 1794±Δ mm against the direction of flight, where ΔX is a value not exceeding the value of 50 mm.

In addition, the length of the anterior compartment in the inventive solution, the aircraft may be selected from a range of 5400 5450...mm, and the rear of the range 5970...6020 mm

In addition, the length of the anterior compartment in the present decision of the Board b is to be selected from a range of 7400 7450...mm, and the rear of the range 7470...7520 mm

The goal of the project claimed the decision "Swept wing" is solved as follows.

Know the swept wing of the plane containing the console and centreplane part. Each of the outer wings are formed by sections of surfaces of single and double curvature smoothly mated with five aerodynamic profiles. Airfoils are placed in the base sections, parallel to the plane of symmetry of the aircraft. The shape of each profile is made up of upper and lower procontrol, and the first plane of the aerofoil is placed in the on-Board cross-section, and the front edges of each subsequent profile is shifted against the direction of flight relative to the leading edge of the previous profile.

The claimed technical solution of the swept wing news is that the cantilever portion of the wing is made of sequentially placed along poluraspada wing, starting from the side section, the first section of the surface of single curvature, the first portion of the surface of double curvature, the second portion of the surface of single curvature and the second surface of double curvature. The first portion of the surface of single curvature is limited to the first and second aerodynamic profile, the first portion of the surface of the double krivi the us is limited to the second and third aerodynamic profiles, the second portion of the surface of single curvature is limited to the third and fourth aerodynamic profile, a second portion of the surface of double curvature is limited to the fourth aerodynamic profile and the wing tip. Each of the surfaces of single curvature formed by segments connecting the contours of the edge of an aerofoil, with their ends placed in the axis of the contours of the same name percent projection profile construction the horizontal fuselage.

In the inventive solution, the plane of the first aerodynamic profile is offset from the plane of symmetry of the aircraft at 1640±ΔZ mm, and its contour defined by the parameters given in table 9.

Table 9
X1YB1YH1
000
0,250,883-0,806
0,51,246-1,149
11,758-1,641
32,970 -2,942
53,683-3,934
7,54,262-5,010
104,635-5,954
155,030-7,530
205,150-8,760
255,103-9,673
304,948-10,312
354,703-10,715
404,366-10,917
453,930-10,946
503,378-10,807
601,964-10,041
700,322-8,792
80-1,460-7,450
90-3,476-6,380
95-4,513-6,003
100-5,540-5,725

The plane of the second aerodynamic profile is offset from the plane of symmetry of the aircraft at 3500±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 1160±Δ mm, and its contour defined by the parameters given in table 10.

Table 10
X2YB2YH2
000
0,250,716-0,701
0,50,981-0,979
11,348-1,378
32,152-2,481
5 2,637-3,279
7,53,071-4,080
103,395-4,749
153,829-5,827
204,073-6,662
254,179-7,339
304,173-7,910
354,076-8,383
403,895-8,730
453,638-8,898
503,313-8,844
602,479-8,083
701,345-6,712
80-0,133-5,148
90-1,920 -4,013
95-2,896-3,851
100-3,877-4,113

The third plane of the airfoil is offset from the plane of symmetry of the aircraft at 5235±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 2245±Δ mm, and its contour defined by the parameters given in table 11.

Table 11
X3YB3VH3
000
0,250,593-0,532
0,50,845-0,738
11,225-1,037
32,120-1,837
52,676-2,450
7,53,175-3,120
103,553-3,713
154,100-4,745
204,468-5,579
254,715-6,231
304,873-6,693
354,961-6,955
404,997-7,027
454,982-6,906
504,910-6,600
604,523-5,416
703,697-3,644
802,404-1,827
900,696-0,879
95-0,257-0,961
100-1,191-1,508

The fourth plane of the airfoil is offset from the plane of symmetry of the aircraft at 11200±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 5340±Δ mm, and its contour defined by the parameters given in table 12.

Table 12
X4YB4YH4
000
0,250,755-0,771
0,51,017-1,033
11,373-1,394
32,204-2,154
52,740-2,604
7,5 3,227-3,027
103,617-3,356
154,212-3,876
204,653-4,303
254,996-4,657
305,260-4,918
355,449-5,057
405,570-5,041
455,626-4,853
505,610-4,505
605,287-3,506
704,558-2,298
803,408-0,955
901,813-0,089
95 0,880-0,208
100-0,057-0,726

The first portion of the surface of double curvature executed smoothly mated with the first and second parts of the surface of single curvature, and the second portion of the surface of double curvature smoothly mated with the second portion of the surface of single curvature and the fifth aerodynamic profile. The fifth plane of the airfoil is offset from the plane of symmetry of the aircraft at 13240±ΔZ mm, the front edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 6410±Δ mm, and its contour defined by the parameters given in table 13.

Table 13
X5YB5YH5
000
0,250,642-0,735
0,50,919-0,998
11,301 -1,353
32,228-2,115
52,825-2,543
7,53,396-2,914
103,834-3,183
154,500-3,549
204,992-3,789
255,382-3,949
305,703-4,022
355,945-4,023
406,111-3,924
456,197-3,704
506,202-3,359
605,934-2,428
705,259 -1,285
804,210-0,068
902,8250,748
952,0300,760
1001,2300,432

In the inventive solution point of the front edges of the first, second, third and fourth aerodynamic profile is shifted from the construction of the horizontal fuselage down on 1010±ΔY mm, 823±ΔY mm, 693±ΔY mm, 103±ΔY mm, respectively, and the point of the leading edge of the fifth airfoil shifted from the construction of the horizontal fuselage up to 95±ΔY mm

In addition, in the inventive solution of the abscissa and ordinates of the points of the upper and lower procontrol aerodynamic profile with respect to the axes, missed points from their front edges parallel to the construction of the horizontal fuselage calculated as ratios

x=Xi*Bi/100,

theIn=YB*Bi/100±ΔP,

theN=YHi*Bi/100±ΔP,

where Xi, YBi, YHi- parameters of the aerodynamic contours of the profiles listed in the following tables 9-13, Bi- the length of the projection corresponding to aerodyn the economic profile for building horizontal fuselage, equal to the first, second, third, fourth and fifth profile 5400±Δ mm, 4240±Δ mm, 3155±Δ mm, 1870±Δ mm 1430±Δ mm, i - number of the airfoil. In addition, the ratio ΔZ, ΔX, ΔY, ΔP, Δ - values, values which do not exceed 100, 30, 40, 5 and 30 mm, respectively.

The technical result from the use of the proposed technical solution of the plane is the ability to design and manufacture on the basis of the model range of passenger regional aircraft, designed to carry from 60 to 110 passengers with high feasibility and technical-operational parameters.

So the sum of the parameters of the nose compartment of the fuselage are shown in table 1, allows not only to improve the aerodynamic characteristics of the aircraft as a whole, but also to provide comfortable accommodation for pilots in the cockpit and provide a good overview of the pilots the runway.

The set of features front, center and rear compartments are given in table 2, 3, 4, 5, 6, allows to provide the equivalent diameter of the fuselage, close to the value of 3.54 m, which is typical for regional passenger aircraft, the possibility of accommodation in the passenger compartment dual power seats on one of the boards and triple blocks on the other side. The implementation of the cross is ecene fuselage in accordance with the claimed parameters, shown in tables 2-6, provides a high level of comfort for passengers: zone diameter to accommodate the head of a passenger exceeds 300 mm, which provides comfortable accommodation of passengers in the seats adjacent to the inner lining of the passenger compartment. The selected dimensions allow to provide the desired height of the passenger compartment along the Central aisle, and the height of the cargo compartment, corresponding to modern requirements.

In addition, the choice of parameters nose, front, center, rear and tail sections, are given in tables 1-7, allows to minimize the length of the fuselage, which allows to reduce the area of wetted surface, to reduce the mass of the structure by reducing the length of wires and pipes.

The configuration of the cross-section of the fuselage with an effective diameter of 3.54 m, providing comfortable accommodation with five chairs in a row, makes it possible to solve the problem of the development of the fuselage of the model aircraft, designed to carry from 60 to 100 passengers. The length of the fuselage can vary from 24 to 29 m, the change in the lengthening of the fuselage λ is in the range from 6.8 to 8.4, which in combination with the claimed features of the relative position of the fuselage and wing allows you to design all the aircraft model range is regionalnych passenger aircraft using wing with the same geometrical characteristics.

The implementation of the swept wing, including the area of double curvature, in combination with placement of the underneath jet engines, can reduce harmful interference between the nacelle, pylon, wing and fuselage, which in combination with the stated parameters of the wing allows the use of one wing geometry for aircraft across the range. In addition, the presence of areas with a single curvature on consoles wing allows to simplify the manufacturing technology wing in production.

In the inventive solution of the plane is most expedient to use the swept wing, the signs of which are disclosed in the claimed technical solution "Swept wing", although in the present plane can be used and wing, but with different parameters.

The vertical and horizontal tail unit in accordance with the stated parameters are shown in table 8, in combination with the parameters of the fuselage and the wing can be placed on them rudders and stabilizers, the effectiveness of which is sufficient for aircraft control across the range. While performing vertical and horizontal tail for the most part in the form of surfaces of single curvature allows to simplify the technology of their production.

Design of the study show that h is about selecting the declared lengths of the front and rear compartments on the basis of the claimed solution can be designed lineup of passenger aircraft, including, for example, the aircraft is designed for transportation from 66 to 83 passengers and from 93 to 108 passengers.

The technical result from the use of the proposed technical solution of the swept wing is the development of the swept wing with high aerodynamic characteristics, allowing its use without dimensional changes in the structure of the model range of passenger aircraft of various passenger capacity and range.

Calculations and tests have shown that the use of the claimed solution, it is possible to develop a swept wing span of 27...28 m, elongation λ=9...10, constriction, no less than ή=3,2. In combination with the nacelle engines positioned below the first surface with double curvature of the swept wing provides high values of aerodynamic quality level Tomax=15...17 in a wide range of Mach numbers, the maximum value of the index Kmax*M on level 12...13 is achieved at M=0.76 to...of 0.82, indicating a high aerodynamic characteristics of the inventive surface.

In addition, the aerodynamic characteristics of the wing and nacelle positioned below the first surface of double curvature, remain stable when you change the mass and length of the fuselage of the aircraft: Aero is DINAMIChESKOE quality of the wing varies by no more than 1%...4% when moving from the aircraft, designed to carry 75 passengers to the aircraft, designed to carry 95 passengers. This allows the use of the inventive wing without changing the geometrical parameters for the model range of passenger aircraft of different Seating capacity.

High aerodynamic characteristics of the claimed swept wing is provided by the choice of the geometric dimensions of the aerodynamic profile, the parameters are given in tables 9-13, and their arrangement relative to each other along the outer wings. This thickened nature of the first aerodynamic profile provides a favorable interference with the fuselage and increases the amount of fuel available in the caisson. The nature of the second and third aerodynamic profiles and their mutual arrangement provides a favorable interference of wing and nacelle engines placed under the wing. The fourth and fifth airfoils form wrap cantilever wing, virtually ensuring the straightness of the Isobar at cruise flight mode.

In addition, the use in aerodynamic design of the wing two limited length along poluraspada consoles surfaces of double curvature simplifies the manufacture of the outer wings in production.

The invention is illustrated by the following drawings:

figure 1 - General view of the proposed airplane;

figure 2 - scheme of division of the proposed airplane;

figure 3 - view of the fuselage side;

4 is a diagram of the nose compartment of the fuselage;

5 is a diagram of the anterior compartment of the fuselage;

6 is a diagram of the Central compartment of the fuselage;

Fig.7. diagram of the wing, connected to the Central compartment;

Fig diagram of the rear compartment of the fuselage;

figure 9 - diagram of the caudal compartment of the fuselage;

figure 10 - diagram of the formation areas of the surfaces of single curvature vertical and horizontal tail;

11 is a view of the tail section from the top (view C with Fig.9);

Fig, 13, 14, 15 is a layout diagram of the aircraft at 95 and 75 passengers;

Fig - General scheme of the swept wing, the view in the plan;

Fig - General scheme of the aerodynamic profile in the lateral projection;

Fig - circuit formation surface of single curvature;

Fig first airfoil;

Fig second airfoil;

Fig - third airfoil;

Fig - fourth airfoil;

Fig - fifth airfoil;

Fig settlement lines wrapping the upper surface of the wing;

Fig - calculated distribution of isobars on the upper surface of the wing;

Fig - dependence of Kmax and M*Kmax from the Mach number;

Fig - dependence of this model the quality of the wing from the lift coefficient at different passenger capacity of the aircraft.

The claimed plane is arranged as follows.

The plane contains (see figure 1, 2) the fuselage 1, the swept wing 2, 3 vertical and 4 horizontal tail surfaces. The fuselage 1 is composed of sequentially placed nasal 5, front 6, Central 7, 8 and the rear tail 9 compartments. Additionally, in the forward compartment 5 can be allocated nose cone 10, and in the tail compartment - compartment for auxiliary propulsion system 11.

Console swept wing for the most part connected to the Central compartment 7 of the fuselage. To improve the aerodynamic characteristics of the aircraft and placement systems and units aircraft supplied ventral Radome 45.

The horizontal stabilizer 4 is connected with the tail compartment. Vertical tail 3 for the most part connected with the tail compartment of the fuselage.

Part of the vertical fin 3 is the rudder 12 and horizontal tail - rudder 13. Two jet engine placed in motomondiale 14 and mounted on pylons 15 podkosami wing 2. Swept wing made from the center section 16, structurally included in the Central compartment 7 of the fuselage and outer wing consoles, which is connected with the center section on the side section 17. To improve the aerodynamic characteristics of the aircraft equipped with salisa wing 18, smoothly paired with polfus legnum Radome 45.

The claimed solution of the plane is set coordinates of a set of points of the outer surface of the skin of the fuselage sections in a rectangular coordinate system. The x-axis compartments combined with building 19 horizontal fuselage lying in the plane of symmetry of the aircraft. The direction of the x-axis for each compartment of the fuselage and parts of some of the compartments are discussed below. The y-axis of each of the compartments located in the plane of symmetry of the aircraft and the axis of applicat compartments perpendicular to the plane of symmetry of the aircraft, and its positive direction for all fuselage sections and parts thereof directed toward the left side of the plane.

The set of points of the outer surface of the fuselage skin selected from the contours formed by sections of the fuselage transverse planes. Each of these contours are symmetric about the plane of symmetry of the aircraft, so each of the paths is determined by points of the half circuit of the left side of the fuselage. Each path is specified by seven points, as the first and second points taken as the point of intersection of the path with the plane of symmetry of the aircraft, as a third point taken contour point, most remote from the plane of symmetry of the aircraft, as the fourth and fifth points taken points on the path, away from the plane of symmetry of the aircraft by half, and the pole is I and the seventh point on the seven-eighths of the distance between the third point and the plane of symmetry of the aircraft. The choice of the number of points in each circuit and the number of paths for each of the compartments and their parts accurately describe the claimed surface of the outer skin of the fuselage.

The outer surface of the lining of the nose compartment of the fuselage is specified in the coordinate system (see figure 4), the abscissa (X11) which is directed against the flight, and the origin of coordinates is placed in the transverse plane containing the tip of the nose of the aircraft. The coordinates of the points of the outer surface of the nose compartment defined by parameters listed above in table 1.

The coordinates of the abscissa and ordinate of the points of each contour, the outer surface of the lining of the nose and other fuselage sections and their parts are calculated based on the following General proportions:

xij=Xij±Δ,

yijk=Yijk±Δ,

and applicati third points of the contours ratio

zij3=Zij3±Δz.

Applicati fourth and fifth points of the contours is equal to half of applicat third points, and applicati sixth and seventh points equal to seven-eighths of applicat third points of the contours.

In the relations Xij, Yijk, Zij3- parameters of the coordinates of the outer surface of the fuselage skin, are given in tables 1-7, i - the index of the compartment, with i=1 for the nose compartment, i=2 for the PE the front compartment, i=3 for the Central compartment, i=4 for the rear compartment, i=5 for the tail compartment. The index j in the proportions indicated by the selected direction of x-axis in the coordinate system of compartments and their parts: j=1, if the direction of the X-axis is chosen against the direction of flight, j=2, if the direction of the X-axis is chosen along the direction of flight. In addition, k is the number of points of the contour (k=1, 2, ...7), and Δ, Δ, ∆ - values not exceeding 5 mm, So the coordinates of the points of each contour, the outer surface of the cladding nose compartment is calculated by the ratio

x11=X11±Δx,

y11k=Y11k±Δ,

z113=Z113±Δz,

where X11, Y11k, Z113the parameters given in table 1.

As in the near equivalent, the length of the front 6 and rear 8 compartments is selected length for a given passenger capacity of the aircraft. However, if in the nearest analogue and a number of other solutions these compartments are cylindrical or cylindrophinae shape with a constant cross-section throughout the length of the front and rear compartments, the claimed solution to the plane of the junction 20 of the nose and front compartments and the plane of junction 23 of the back and tail compartments placed on parts of the fuselage, tapering to the nose of the plane and the tip of the fuselage. In the inventive solution plosko and interface 21 and 22 of the front 6 and rear 8 compartments Central compartment 7 is also placed on the parts of the fuselage with the changing configuration along the longitudinal axis of the fuselage contour of the cross section. This provides the optimum combination of technological design of the fuselage in its manufacture with high aerodynamic characteristics of the aircraft. This causes, on the other hand, the special assignment of parts of the front compartment, adjacent to the nasal and front fuselage sections and parts of the rear compartment adjacent to the main and tail fuselage sections are shown in tables 2, 3, 5, 6.

Part 24 of the front surface of the compartment 6 of the fuselage adjacent to the junction of the front compartment with the nose compartment 5, is set in the coordinate system (see figure 5), the abscissa (X21) which is directed against the flight, and the origin of coordinates is placed in the transverse plane, combined with the plane of junction 20 of the nose and front compartments, the value of the parameter coordinates of surface points of the specified portion of the nose compartment are shown in table 2, and their coordinates are calculated by the above ratios.

Part 25 of the front surface of the compartment adjacent to the Central compartment, specified in the coordinate system, the x-axis (X22) which points in the direction of flight, origin combined with the plane of junction 21 of the front compartment to the Central compartment, and the parameters of coordinates of points on the surface of this part of the front compartment are shown in table 3, and their coordinates are calculated for given the mentioned above ratios.

Between the parts 24 and 25 of the front compartment contains a part of the front cover 26 with a constant along the length of the contour shape of the cross section that allows without changing the shape of the parts 24 and 25 to change its length to obtain the required passenger capacity of the aircraft.

The outer surface of the cladding of the Central compartment of the fuselage is specified in the coordinate system (see Fig.6), the abscissa (X31) which is directed against the flight, and the origin of coordinates aligned with the plane of junction 21 of the Central compartment to the front compartment, while the value of the parameters the coordinates of surface points of the specified portion of the nose compartment is shown in table 4, and their coordinates are calculated by the above ratios.

As mentioned above, the fuselage of the aircraft is connected swept wing. The outer surface of the cladding of the cantilever part of the swept wing made of sequentially placed along poluraspada wing, starting from the side section 17 (see Fig.7), is smoothly conjugate to each other in the first area of the surface of single curvature 29, the first portion of the surface of double curvature 30, the second portion of the surface of single curvature 31 and the second portion of the surface of double curvature 32. In addition, the outer surface of the consoles of the swept wing smoothly paired with aerodynamic profiles, one of which 27 RA is substituted in the side section 17 of the fuselage, offset from the plane of symmetry of the plane at z01=1640±ΔZ mm. In contrast to the nearest analogue of the wing for the most part connected with the Central compartment 7 of the fuselage, but the front part of the specified aerodynamic profile 27 placed in the on-Board section 17, is connected with the front compartment 6 of the fuselage. At this point its front edge 28 is displaced forward in the direction of flight from the Central junction 7 and the front 6 of the fuselage sections at X'1Co=480±Δ mm, where ΔZ and ΔX is a value, not exceeding 100 and 30 mm, respectively.

Furthermore, the said jet engines in the inventive solution is placed under the first sections 30 double curvature consoles swept wing. These features accommodate the swept wing in combination with the geometrical characteristics of the fuselage and swept wing, disclosed below, provide the most optimal aerodynamic characteristics of a passenger plane.

Part 33 of the rear surface of the compartment 8 of the fuselage, adjacent to the junction 22 of the rear compartment 8 with the Central 7, specified in the coordinate system (see Fig), the abscissa (X41) which is directed against the flight, and the origin of coordinates is placed in the transverse plane, combined with the plane of junction 22 of the Central and rear compartments, the value of the parameters of the coordinate points of the outer is poverhnosti specified portion of the rear compartment are shown in table 5, and their coordinates are calculated by the above ratios.

The portion 34 of the surface of the rear cover 8 adjacent to the tail compartment 9, is set in the coordinate system (see Fig), the abscissa (X42) which points in the direction of flight, origin combined with the plane of junction 23 of the rear compartment with the tail compartment, and the parameters of coordinates of points on the surface of this part of the rear compartment are shown in table 6, and their coordinates are calculated by the above ratios.

Between the parts 33 and 34 of the rear compartment and placed part of the rear compartment 35 with a constant along the length of the contour shape of the cross section that allows without changing the shape of the parts 33 and 34 to change its length to obtain the required passenger capacity of the aircraft.

The outer surface of the shell of the tail cover 9 (see Fig.9) of the fuselage is specified in the coordinate system, the x-axis (X51) which is directed against the flight, and the origin of coordinates aligned with the plane of junction 23 of the tail and rear compartments. The parameters of the coordinates of surface points of the tail compartment are shown in table 7, and their coordinates are calculated by the above ratios.

As mentioned above, the fuselage is connected horizontal and vertical tail surfaces. The most appropriate vertical and horizontal tail to run for more the x part in the form of plots of the surfaces of single curvature. Sections of surfaces of single curvature vertical and horizontal tail when it is formed by two symmetric boundary aerodynamic profiles of the first 36 and second 37 (see figure 10) and segments 38, connecting the contours of the edge of an aerofoil. The ends of the segments are placed in the y-paths of the same name percent of the chord of an aerofoil.

The first plane of the airfoil 38 vertical plane, it is advisable to shift from the construction of the horizontal fuselage on the1in=1540±ΔY mm, and the second 392in=6810±ΔY (see Fig.9). The plane of the first aerodynamic profile of the horizontal tail is displaced in the z1th=1100±ΔZ mm from the plane of symmetry of the aircraft, and the second at z2th=5170±ΔZ mm (see 11). The point of the leading edge 42 of the first airfoil 40 horizontal tail is shifted upward on the1th=750±ΔY mm from the construction of the horizontal fuselage of the plane, with the horizontal stabilizer is made with an angle transverse V on the front edge of 3.5°...4,5° and sweep 34...35 degrees, and the vertical tail with the sweep of the leading edge 39-40 degrees. The contours of the first and second aerodynamic profiles of the vertical and horizontal tail set parameters are shown in table 8. PR is this the abscissa and ordinate of the contours of the aerodynamic profiles of the vertical and horizontal tail are calculated from ratios

x=X0*/100,

y=Y0*B/100±ΔP,

where X0, Y0- parameters of the aerodynamic contours of the profiles listed in the following table 8, the length of the corresponding chord of the airfoil, equal to the first and second aerodynamic profiles of the vertical fin 4496±Δ mm and 1364±Δ mm, respectively, and first and second aerodynamic profiles of the horizontal tail 2548±Δ mm and 815±Δ mm, respectively, while in the above ratios Δ, ΔPΔYP, ΔY, ΔZ - value not exceeding 30, 30, 5 and 40 mm, respectively.

The point of the leading edge 43 of the first airfoil 38 vertical plane it is advisable to move forward in the flight direction from the plane of the junction 23 of the tail and rear compartments for x1in=951±ΔX mm, and the point of the leading edge 42 of the first airfoil horizontal tail - x1th=1794±Δ mm, where Δ - value not exceeding the value of 50 mm.

Such geometrical parameters of vertical and horizontal tail enable the use of the claimed solution in an aircraft designed for regional transportation, be placed on the tail plumage of the Executive bodies (wheels) with geometric dimensions, providing effective management of the aircraft when cheating the AI its Seating capacity without changing the geometrical parameters of the Executive bodies.

The inventive configuration of the transverse shape of the fuselage of a passenger aircraft can accommodate inside the passenger compartment of five seats in a row with the required comfort conditions for passengers. The use of the claimed solution, the length of the front compartment can be selected from the range of L6=5400 5450...mm (see Fig), and the rear of the range L8=5970...6020 mm in the passenger cabin can be placed from 66 to 83. The use of the claimed solution, the length of the front compartment can also be selected from the range of L6=7400 7450...mm, and the rear of the range L8=7470...7520 mm in the passenger cabin can be placed from 93 to 108 passengers. Layout schemes of these solutions is shown in Fig-16.

Obtained in accordance with the proposed technical solution section of the fuselage allows you to put in the inner space of the passenger compartment of the double blocks of seats on one of the boards and triple blocks of seats on the other side, as shown in Fig. The distance between blocks of seats is about 508 mm, which allows easy separation of the two bogies for passenger service. The height from the floor to the ceiling of the passenger compartment can be SPW=210 mm (see Fig), which is sufficient for the passage of passengers. In addition, the thickness of the as the design of the walls of the fuselage, including trim, power set and the inner panel, which in a typical performance does not exceed 100 mm, provides comfortable accommodation of a passenger in the extreme to a side chair. (See Fig) sphere radius to accommodate the head of a passenger is about 300 mm, which corresponds to modern requirements for comfortable accommodation of passengers. Level of comfort can be improved with the placement on each side of the dual blocks of seats.

In addition, the claimed technical solution of the fuselage provides the height of the cargo compartment of Nbgo=1016 mm, which corresponds to modern requirements for their height.

Accommodation with a high level of comfort five seats in a row allows you to use the claimed technical solution of the fuselage in the family of regional aircraft various passenger preserving the transverse configuration of the fuselage and the use of one wing geometry. The design of the inventors show that the proposed configuration of the fuselage does not require changing the geometry of the wing by varying the length of the fuselage from 24 to 30 feet, It provides elaboration on the basis of this technical solution family of regional aircraft, designed, for example, to travel from 60 to 100 passengers, with a high degree of unification.

The most appropriate use is with in the present decision of the passenger aircraft disclosed below, the decision of the swept wing.

Declare swept wing is arranged as follows.

Console part 51 (7, 16) of the swept wing of the plane is made of sequentially along poluraspada wing of the first portion of the surface of single curvature 29, the first portion of the surface of double curvature 30, the second portion of the surface of single curvature 31 and the second portion of the surface of double curvature 32.

Each of the parts of the surfaces of single curvature 29 and 31, as shown in Fig, formed by segments 56. Segments 56 connect circuits 57 aerodynamic profiles, available at boundaries between areas of the surface of the wing. Thus the abscissa of these 57 points equidistant from the front edge 63 of the profile along the axis 58, missing from the front edge profiles parallel to the construction of the horizontal fuselage.

The surface of the single and double curvature of the proposed solutions seamlessly paired with five aerodynamic profiles 27, 52, 53, 54, 55, located in the base sections, parallel to the plane of symmetry of the aircraft. The contour of each of the profiles compiled (see Fig) from the top 59 and the bottom 60 procontrol. Procentury are connected at the points of the leading edge 63 of the wing and tail profile. Thus the top and bottom procentury profiles are defined by the set of coordinates of their points of abscissa of which are set along the OS is, missed the point of the leading edge profile parallel to the construction of the horizontal fuselage, and the ordinate yinand In(see 17) in the perpendicular direction.

Airfoils declare the swept wing in the planes of the base sections is specified by the coordinates of the contours of the profiles (see Fig). The General scheme describing the aerodynamic profiles 27 and 52-55 declare the swept wing is shown in Fig. The main parameters of the profile are: Bi - length projection profile construction the horizontal fuselage and the parameters of the coordinate points of the contours of the airfoil in the coordinate system XOY in the plane of the base section.

Abscissas and ordinates of the points of the upper and lower procontrol aerodynamic profile with respect to the axes, missed points from their front edges parallel to the construction of the horizontal fuselage, are calculated by the relations

x=Xi*Ini,

yB=YiB*Ini±ΔP,

yH=YiH*Bi±ΔP,

where X0, YB0, YH0- parameters of the aerodynamic contours of the profiles listed in the following tables 9-13.

The length of the projection of the first, second, third, fourth and fifth aerodynamic profiles for building horizontal fuselage in the inventive solution streamign the wing is equal to the first, the second, third, fourth and fifth profile B1=5400±Δ mm, In2=4240±Δ mm, In3=3155±Δ mm, In4=1870±Δ mm, In5=1430±Δ mm In the ratios of X0, YB0, YH0- parameters of the aerodynamic profiles are shown in tables 9-13 and values ΔP, Δ not exceed the values 5 and 30 mm, respectively.

The plane of the first profile 27 is placed in the on-Board cross-section and offset from the plane of symmetry of the plane at z01=1640±ΔZ mm Plane 52 of the second, third 53 and fourth 54 and 55 fifth aerodynamic profile is offset from the plane of symmetry of the plane at z02=3500±ΔZ mm, z03=5235±ΔZ mm, z04=11200±ΔZ mm, z05=13240±ΔZ mm, the Value of ΔZ is less than 100 mm, While the aerodynamic characteristics of the wing are virtually unchanged.

In the planes of the base sections of the point 63 of the front edges of each of the following aerodynamic profiles (see Fig, 19-23) are displaced relative to each other in the flight direction and in the vertical direction.

In the direction of flight of the point of the front edge 63 of the aerodynamic profile is shifted against the direction of flight relative to the point of the leading edge of the previous profile. Point 63 of the front edges 52 of the second, third 53 and fourth 54 and 55 fifth aerodynamic profiles shifted relative to the point 28 front the second edge of the first airfoil at x 02=1160±Δ mm, x03=2245±Δ mm, x04=5340±Δ mm and x05=6410±Δ mm, respectively. Value Δ not exceed 30 mm In Fig-23 offset points 63 of the front contour of the profile in the direction of flight indicated as x0iwhere i is the number of the aerodynamic profile.

In the vertical direction (see Fig-23) point of the leading edge 28 of the first aerodynamic profile 27 and the point 63 of the front edges 52 of the second, third 53 and fourth 54 aerodynamic profile is shifted from the construction of the horizontal fuselage down on y01=1010±ΔY mm, y02=823±ΔY mm, y03=693±ΔY mm, y04=103±ΔY mm, respectively, and the point 63 of the leading edge of the fifth airfoil shifted from the construction of the horizontal fuselage up on the05=95±ΔY mm, the Value of ΔY is less than 40 mm On Fig-23 offset with 28 points and 63 of the front contour of the profile in the vertical direction is designated as y0iwhere i is the number of the aerodynamic profile.

Given the values Δ and ΔY not lead to significant changes in the aerodynamic characteristics of the wing.

As indicated above mentioned first portion of the surface of single curvature 29, the first portion of the surface of double curvature 30, the second portion of the surface of single curvature 31 and the second section of double curvature 32 sequentially placed is along poluraspada wing, starting from the side section.

The first portion of the surface of single curvature is limited to the first 27 and second 52 aerodynamic profile, the pattern formation of the above.

The first portion of the surface of double curvature limited to 30 second 52 and third 53 aerodynamic profiles. Besides the change in curvature along the direction of flight on the part of the surface curvature changes and longitudinally along the wing direction. In longitudinal along the wing towards the lower surface of this area it is advisable to perform is slightly concave on the inside wing. Under the first portion of the surface of double curvature, it is advisable to place the engine and its nacelle and wing pylon of the engine shown in Fig.7. The first section of double curvature should be performed, in addition smoothly with the paired first and second parts of the surface of single curvature.

The second section 31 of the surface of single curvature is limited to third 53 and fourth 54 aerodynamic profiles.

The second portion of the surface of double curvature is limited to the fourth 54 aerofoil and wing tip 61. As the first section 30 of the surface of double curvature, the second section 32 of the surface of double curvature alters the curvature of the surface in the direction of flight and the direction in which ol wing. The second section of double curvature smoothly mated with the second portion of the surface of single curvature and the fifth 55 aerodynamic profile, the base section which is placed in the inner area of the second segment of double curvature.

The use of the inventive combination of the geometric dimensions of the swept wing allows you to develop a swept wing with a sweep angle of about 25 degrees on the leading edge. When the wing span of 27...28 m wing span is λTr=9,82, narrowing ηTr=3,25. As can be seen (see Fig.7 and 16), the claimed swept wing has a small influx of the leading edge and direct influx trailing edge of the wing on the first surface of single curvature.

Thickened nature of the first aerodynamic profile allows you to put in the wing increased fuel reserves, and around him the air flow reduces the root effect, provides a favorable interference supercritical wing with the fuselage. More forward position of the maximum thickness of the first airfoil compared to the position of maximum thickness along the profile of other aerodynamic profile helps in the flow of the first portion of the surface of single curvature straightening Isobar and improve flow in the basal area of the wing.

The first party is to the surface of double curvature and geometrical parameters of the second and third aerodynamic profile when the flow of incoming flow reduce the impact of harmful interference engine nacelle to the wing.

In addition, the third airfoil, as the initial profile of the second portion of the surface of single curvature, when the flow of incoming flow has the greatest effect on the aerodynamic and weight characteristics of the wing. The third and fourth airfoils, forming a second portion of the surface of single curvature, when the flow of incoming flow provide rigidity Isobar at cruise flight mode.

The geometrical characteristics of the fifth aerodynamic profile, determining the nature of the second portion of the surface of double curvature, provide flow deviation from the elliptical distribution of circulation along the span direction of the bell.

As can be seen from Fig-27, the use of the claimed solution of the swept wing is possible to develop a swept wing span of 27...28 m in combination with the nacelle engines positioned below the first surface with double curvature with high aerodynamic quality level Kmax=15...17 in a wide range of Mach numbers, the maximum value of the index Kmax*M on level 12...13 is achieved at M=0.76 to...of 0.82, indicating a high aerodynamic characteristics of the inventive surface.

In addition, the aerodynamic characteristics of the wing with the engine nacelles under the first participants of the AMI surfaces of double curvature remain stable when you change the mass and length of the fuselage of the aircraft so the aerodynamic performance of the wing varies by no more than 1%...4% when moving from the aircraft, designed to carry 75 passengers, the aircraft is designed to carry 95 passengers. This allows the use of the inventive wing without changing the geometrical parameters for the development of the model range of passenger aircraft of various passenger capacity and range.

The proposed solutions of the plane and the swept wing can be made to the aviation industry.

1. The plane containing the fuselage, United with him swept wing, vertical and horizontal tail, jet engines mounted on pylons under the wing consoles, while the fuselage is made up of sequentially placed bow, front, center, rear and tail sections, characterized in that the outer surface of the fuselage specified coordinates in a rectangular coordinate system, the x-axis which is combined with the construction of the horizontal fuselage, the ordinate axis located in the plane of symmetry of the plane and the axis of applicat perpendicular to the plane of symmetry of the aircraft, and the points of the surface of the fuselage selected from the contours formed by sections of the fuselage transverse planes, each contour is specified by seven points, the first and second points taken as the point of intersection of the path with the plane of symmetry of the aircraft, the quality is as the third point single point circuit, most remote from the plane of symmetry of the aircraft, as the fourth and fifth points taken points on the path, away from the plane of symmetry of the aircraft by half, and the sixth and seventh points on seven-eighths of the distance between the third point and the plane of symmetry of the aircraft, the surface of the nose compartment of the fuselage is specified in the coordinate system, the x-axis which is directed against the flight, and the origin of coordinates is placed in the transverse plane containing the tip of the nose of the plane, with coordinates of points on the surface of the nasal compartment defined by parameters:

td align="center"> -530,0
X11Y111Y112Y113Z113Y114Y115Y116Y117
0-980,0-980,0-980,00,0-980,0-980,0-980,0-980,0
15-885,0-1061,7-980,0107,6-895,9-1050,8-930,2-1019,6
50-801,2-1127,6-979,7197,9-820,7-1107,8-885,0-1051,3
100-719,4-1186,5-978,7281,9-747,0-1158,7-839,7-1079,3
300-495,4-1327,2-968,5496,7-540,6-1279,1-703,9-1142,1
600-248,5-1454,4-936,9714,4-305,1-1385,1-1187,4
1200150,3-1613,0-827,81024,7to 85.2-1507,8-200,6-1207,9
1800650,4-1714,8-679,11242,0508,3-1576,0108,3-1180,5
24001081,9-1785,0-513,01400,0914,9-1614,6spreads for about 319.2-1128,8
30001338,2-1833,5-349,31516,11167,3-1634,6509,5-1067,8
34951481,7-1860,8-228,5 1297,6-1642,1644,8-1018,7

the portion of the surface of the anterior compartment of the fuselage adjacent to the junction of the front compartment with the nose compartment, specified in the coordinate system, the x-axis which is directed against the flight, and the origin of coordinates is placed in the transverse plane, combined with the plane of junction of the nose and front compartments, while the coordinates of the surface points of the specified portion of the nose compartment defined by parameters:
-1881,2
X21Y211Y212Y213Z213Y214Y215Y216Y217
01481,7-1860,8-228,51588,11297,6-1642,1644,8-1018,7
6001601,5-112,11652,71401,1-1644,8742,1-969,4
12001677,3-1889,6-37,51697,01462,1-1646,4792,9-940,2
18001718,2-1890,0-4,91722,51491,4-1646,7821,6-927,0
24001730,0-1890,00,01730,01498,2-1647,5837,5-924,4

and part of the front surface of the compartment adjacent to the Central compartment, specified in the coordinate system, the x-axis which points in the direction of flight, origin combined with the plane of junction of the front compartment to the Central compartment, and the coordinates of the points of the surface the surface of this part of the front compartment defined by parameters:
X22Y221Y222Y223Z223Y224Y225Y226Y227
01730,0-2069,90,01730,01498,2-2062,5837,5-1851,9
6001730,0-2036,70,01730,01498,2-2028,2837,5-1746,3
12001730,0-1998,10,01730,01498,2-1982,3837,5-1460,8
800 1730,0-1957,70,01730,01498,2-1901,1837,5-924,4
24001730,0-1917,80,01730,01498,2-1762,7837,5-924,4
30001730,0-1890,00,01730,01498,2-1647,5837,5-924,4

the surface of the Central compartment of the fuselage is specified in the coordinate system, the x-axis which is directed against flight origin combined with the plane of junction of the Central compartment to the front compartment, and the coordinates of the points of the surface of the Central compartment defined by parameters:
X31Y311Y312Y313 Z313Y314Y315Y316Y317
01730,0-2069,90,01730,01498,2-2062,5837,5-1851,9
6001730,0-2094,50,01730,01498,2-2087,4837,5-1903,7
12001730,0-2109,20,01730,01498,2-2102,2837,5-1932,7
18001730,0-2115,20,01730,01498,2-2108,28375 -1946,0
24001730,0-2113,70,01730,01498,2-2106,8837,5-1944,8
30001730,0-2103,60,01730,01498,2-2096,6837,5-1935,1
36001730,0-2082,70,01730,01498,2-2075,6837,5-1913,8
42001730,0-2052,30,01730,01498,2-2044,9837,5-1878,6
48001730,0-2017,10,01730,0 1498,2-2008,3837,5-1824,4
54001730,0-1981,70,01730,01498,2-1966,6837,5-1738,7
55001730,0-1975,90,01730,01498,2-1958,6837,5-1718,4

the portion of the surface of the rear compartment, adjacent to the Central compartment, specified in the coordinate system, the x-axis which is directed against the direction of flight, and its origin is aligned with the plane of junction of the rear compartment to the Central compartment, while the coordinates of the surface points of this part of the rear compartment defined by parameters:
X41Y411Y412Y413Z413Y414 Y415Y416Y417
01730,0-1975,90,01730,01498,2-1958,6837,5-1718,4
6001730,0-1942,60,01730,01498,2-1894,0837,5-1522,8
12001730,0-1910,60,01730,01498,2-1797,9837,5-924,4
18001730,0-1890,00,01730,01498,2-1679,7837,5-924,4
1902730,0 -1890,00,01730,01498,2-1647,5837,5-924,4

the portion of the surface of the rear compartment of the fuselage, adjacent to the junction of the rear compartment with the tail compartment, specified in the coordinate system, the x-axis which points in the direction of flight, and the origin of coordinates is placed in the transverse plane, combined with the plane of junction of the rear compartment with the tail compartment, while the coordinates of the surface points of this part of the rear compartment defined by parameters:
-1603,2
X42Y421Y422Y423Z423Y424Y425Y426Y427
01699,1-1304,9197,11502,01497,9-1103,6924,2 -530,0
6001721,0-1446,4120,21600,71506,5-1232,7895,2-622,4
12001729,5-1574,656,61672,91505,3-1348,8866,5-701,8
18001730,0-1686,313,61716,31500,0-1452,1844,6-775,6
24001730,0-1777,70,11729,91498,2-1538,7837,5-839,5
30001730,0-1844,60,01730,01498,2837,5-888,6
36001730,0-1882,60,01730,01498,2-1640,2837,5-918,0
39971730,0-18900,01730,01498,2-1647,5837,5-924,4

the surface of the caudal compartment of the fuselage is specified in the coordinate system, the x-axis which is directed against the flight, and the origin of coordinates aligned with the plane of junction of the tail and rear compartments, while the coordinates of the points of the tail compartment defined by parameters:
X51Y511Y512Y513Z513Y514Y515Y516 Y517
01699,1-1304,9197,11502,01497,9-1103,6924,2-530,0
6001665,2-1152,6350,51381,81483,1-963,5976,8-424,3
12001621,4-991,8606,31247,81463,5-819,81062,2-329,5
18001568,9-823,9845,81110,81435,6-670,81140,7-234,5
24001507,4-650,3949,5981,1-514,61169,9-127,8
30001439,2-473,8945,4859,01355,9-353,51157,5-10,7
36001370,8-297,2940,3736,91306,6-192,31138,0106,6
42001302,4-120,6935,1614,81252,1-31,11111,0223,9
48001233,856,1928,5491,91193,0130,21076,4341,4
54001125,9253,8 869,2341,21091,8309,6994,3467,6
5874,5909,9523,2734,5175,3886,8548,9820,9622,6

thus the abscissa and ordinates of the points of the contours of the compartments is calculated according to the following ratios:
xij=Xij±Δx,
yijk=Yijk±Δy,
and applicati third points of the contours ratio:
zij3=Zij3±Δz,
in which Xij, Yijk, Zij3- the specified parameters of the coordinate points of the paths, i is the index of the compartment, with i=1 nasal compartment, i=2 for the front compartment, i=3 for the Central compartment, i=4 for the rear compartment, i=5 for the tail compartment, j=1 for the nose compartment, front compartment, adjacent to the nose compartment of the rear compartment, adjacent to the Central compartment, and the tail compartment, j=2 for the front part of the compartment adjacent to the Central compartment and the rear compartment adjacent to the tail compartment, k - the number of points of the contour, a Δx, Δy, Δz - values not exceeding 5 mm, the surface of the cantilever frequent is referred to the swept wing made of sequentially placed along poluraspada wing, starting from the side section is smoothly conjugate to each other in the first area of the surface of single curvature, the first portion of the surface of double curvature, the second portion of the surface of single curvature and the second surface of double curvature, in addition, the surface of the cantilever part of the wing smoothly paired with aerodynamic profiles, one of which is placed in the on-Board cross-section of the fuselage, offset from the plane of symmetry of the aircraft at 1640±ΔZ mm, and its leading edge is offset from the junction of the Central and anterior compartments of the fuselage forward flight at 480±Δ mm, where ΔZ and ΔX is a value, not exceeding 100 and 30 mm, respectively, and referred to jet engines placed under the first sections of double curvature consoles swept wing.

2. The aircraft according to claim 1, characterized in that the horizontal and vertical plumage on their large pieces made in the form of plots of single curvature, the surface of which is formed by two boundary symmetrical aerodynamic profile and line segments connecting the contours of the edge of an aerofoil, and the ends of the segments are placed in the y-paths of the same name percent of the chord of an aerofoil, with the plane of the first vertical tail is shifted upward from construction l the fuselage at 1540±ΔY mm, and second - 6810±ΔY mm, the plane of the first airfoil horizontal tail shifted 1100±ΔZ mm from the plane of symmetry of the aircraft, and the second - 5170±ΔZ mm, and the point of the front edge of the first airfoil horizontal tail shifted up to 750±ΔY mm from the construction of the horizontal fuselage, horizontal tail is made with an angle transverse V on the front edge of 3.5°...4,5° and sweep 34...35°, and the vertical tail with the sweep of the leading edge 39-40°, and the contour of the first and second aerodynamic profiles and vertical horizontal tail set parameters:

X0Y0
0,00,0
0,250,697
0,50,987
11,373
32,236
52,770
7,53,256
103,631
15 4,191
254,823
304,957
354,991
404,931
504,530
603,844
702,972
801,998
900,992
1000,0

thus the abscissa and ordinate of the contours of the aerodynamic profiles of the vertical and horizontal tail is calculated according to the following ratios:
x=X0·/100,
y=Y0·/100±ΔP,
where X0, Y0- parameters of the aerodynamic contours of the profiles listed in the following table, the length of the corresponding chord of the airfoil, equal to the first and second aerodynamic profile of the vertical fin 4496±Δ mm and 1364±Δ mm, respectively, and for the first and second airfoil horizontal tail 2548±Δ mm and 815±Δ mm, respectively, while in the preceding ratios, ΔY, ΔZ, ±ΔPΔ, - is elicina, not exceeding 30, 30, 5,40 mm, respectively.

3. The aircraft according to claim 2, characterized in that the toe of the first airfoil vertical plane offset from the plane of the junction of the tail and rear compartments 951±Δ mm forward direction of flight, and the toe of the first airfoil horizontal tail - 1794±Δ mm against the direction of flight, where ΔX is a value not exceeding the value of 50 mm.

4. The aircraft according to claim 3, characterized in that the length of the anterior compartment selected from the range of 5400 5450...mm, and the rear of the range 5970...6020 mm

5. The aircraft according to claim 3, characterized in that the length of the anterior compartment selected from a range of 7400 7450...mm, and the rear of the range 7470...7520 mm

6. The swept wing of the plane containing the console and centreplane part, each of the outer wings are formed by sections of surfaces of single and double curvature smoothly mated with five aerodynamic profiles and placed in the base sections, parallel to the plane of symmetry of the aircraft, and the shape of each profile is composed of upper and lower procontrol, with the first plane of the profile is placed in the on-Board cross-section and point of the front edges of each subsequent profile is shifted against the direction of flight relative to the point of the leading edge of the previous profile, characterized in that the con is Aulnay part of the wing is made of sequentially placed along poluraspada wing, starting from the side section, the first section of the surface of single curvature, which is limited to the first and second aerodynamic profile, the first portion of the surface of double curvature, which is limited to the second and third aerodynamic profile, the second portion of the surface of single curvature, limited third and fourth aerodynamic profile, and the second portion of the surface of double curvature bounded fourth aerodynamic profile and the wing tip, where each of the surfaces of single curvature formed by segments connecting the contours of the edge of an aerofoil, with their ends placed in the axis of the contours of the same name percent projection profile construction the horizontal fuselage, and the plane of the first aerodynamic profile is offset from the plane symmetry plane at 1640±ΔZ mm, and its contour defined parameters:

X1VB1YH1
000
0,250,883-0,806
0,51,46 -1,149
11,758-1,641
32,970-2,942
53,683-3,934
7,54,262-5,010
104,635-5,954
155,030-7,530
205,150-8,760
255,103-9,673
304,948-10,312
354,703-10,715
404,366-10,917
453,930-10,946
503,378-10,807
601,964-10,041
700,322-8,792
80-1,460-7,450
90-3,476-6,380
95-4,513-6,003
100-5,540-5,725

the plane of the second aerodynamic profile is offset from the plane of symmetry of the aircraft at 3500±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 1160±Δ mm, and its contour defined parameters:
X2YB2YH2
000
0,250,716-0,701
0,50,981-0,979
11,348-1,378
32,152-2,481
52,637-3,279
7,53,071-4,080
103,395-4,749
153,829-5,827
204,073-6,662
254,179-7,339
304,173-7,910
354,076-8,383
403,895-8,730
453,638-8,898
503,313-8,844
602,479 -8,083
701,345-6,712
80-0,133-5,148
90-1,920-4,013
95-2,896-3,851
100-3,877-4,113

the third plane of the airfoil is offset from the plane of symmetry of the aircraft at 5235±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 2245±ΔY mm, and its contour defined parameters:
X3YB3YH3
000
0,250,593-0,532
0,50,845-0,738
11,225 -1,037
32,120-1,837
52,676-2,450
7,53,175-3,120
103,553-3,713
154,100-4,745
204,468-5,579
254,715-6,231
304,873-6,693
354,961-6,955
404,997-7,027
454,982-6,906
504,910-6,600
604,523-5,416
70 3,697-3,644
802,404-1,827
900,696-0,879
95-0,257-0,961
100-1,191-1,508

the fourth plane of the airfoil is offset from the plane of symmetry of the aircraft at 11200±ΔZ mm, the point of its leading edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 5340±Δ mm, and its contour defined parameters:
X4YB4YH4
000
0,250,755-0,771
0,51,017-1,033
11,373-1,394
3 2,204-2,154
52,740-2,604
7,53,227-3,027
103,617-3,356
154,212-3,876
204,653-4,303
254,996-4,657
305,260-4,918
355,449-5,057
405,570-5,041
455,626-4,853
505,610-4,505
605,287-3,506
704,558-2,298
803,408-0,955
901,813-0,089
950,880-0,208
100-0,057-0,726

the first portion of the surface of double curvature executed smoothly mated with the first and second parts of the surface of single curvature, and the second portion of the surface of double curvature smoothly mated with the second portion of the surface of single curvature and the fifth aerodynamic profile, the plane of which is offset from the plane of symmetry of the aircraft at 13240±ΔZ mm, the front edge is displaced along the direction of flight relative to the point of the front edge of the first airfoil at 6410±Δ mm, and its contour defined parameters:
X5VB5YH5
000
0,250,642-0,735
0,5 0,919-0,998
11,301-1,353
32,228-2,115
52,825-2,543
7,53,396-2,914
103,834-3,183
154,500-3,549
204,992-3,789
255,382-3,949
305,703-4,022
355,945-4,023
406,111-3,924
456,197-3,704
506,202-3,359
605,934-2,428
705,259-1,285
804,210-0,068
902,8250,748
952,0300,760
1001,2300,432

and the point of the front edges of the first, second, third and fourth aerodynamic profile is shifted from the construction of the horizontal fuselage down on 1010±ΔY mm, 823±ΔY mm, 693±ΔY mm, 103±ΔY mm, respectively, and the point of the leading edge of the fifth airfoil shifted from the construction of the horizontal fuselage up to 95±ΔY mm, the abscissa and ordinates of the points of the upper and lower procontrol aerodynamic profile with respect to the axes, missed points from their front edges parallel to the construction of the horizontal fuselage, calculated according to the following ratios:
x=Xi·Bi/100,
thein=YIni·Ini/100±ΔP,
then=YNi·Ini/100±ΔP,
where Xi, YBi, YHi- parameters conturo the aerodynamic profiles listed in the following tables i - number airfoil,i- the length of the projection of the corresponding aerodynamic profile for building horizontal fuselage, equal to the first, second, third, fourth and fifth profile 5400±Δ mm, 4240±Δ mm, 3155±Δ mm, 1870±Δ mm 1430±Δ mm, respectively, a ΔZ, ΔX, ΔY, ΔP, Δ, - size, not exceeding 100, 30, 40, 5, 30 mm, respectively.



 

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Aircraft ramp door // 2389643

FIELD: transport.

SUBSTANCE: invention can be used in aircraft engineering and can be used for boarding/deplaning passengers and crew. Proposed ramp door comprises covered carcass, stairs made from stationary section secured on carcass inner side and rotary section pivoted to stationary section, door opening hydraulic cylinder, folding handrail made up on inner and outer sections and mechanism to open rotary section of stairs. Ramp door carcass is pivoted to airframe with the help of bracket. Aforesaid hydraulic cylinder is secured in door carcass and pivoted to via frame via bell crank. Aforesaid bracket hinge can elevate ramp door relative to airframe. Folding handrail inner section is hinged to airframe and stationary section of stairs. Outer section of folding handrail is coupled with rotary section opening mechanism representing a lever-cable mechanism arranged in ramp door carcass.

EFFECT: simplified design, higher reliability.

2 cl, 7 dwg

FIELD: machine building.

SUBSTANCE: invention refers to packing facilities for sealing automatically closing covers of aircraft hatches. The facility consists of a rubber packing element installed on cover 5 of the hatch. The packing element is made as thin ring 1 secured on cover 5 of the hatch with its cantilever-arranged internal part, while its external part designed to contact packed surface 22 of fringe 15 of the aircraft case is free. The internal part of the ring is clamped along whole perimetre of hatch cover 5 to ensure said cantilever attachment and simultaneously to give a conic shape to the packing element. Hold downs 3 and 4 have congruous conic surfaces contacting surfaces of ring 1 with their tops turned to the side of sealed volume. Hold down 4 is attached by glue to contact surface of the internal part of ring 1. Cover 5 of the hatch and fringe 15 of the aircraft case can be made with slit 21 between them and with pocket 2 along whole perimetre of the external part of the packing element; also pocket 2 communicates with environment through this slit.

EFFECT: invention facilitates pressure tightness of split joint of hatch covers operating under extreme sub and above zero temperatures avoiding application of heavy efforts for closing.

3 cl, 4 dwg

FIELD: transport.

SUBSTANCE: invention relates to aircraft building, particularly to window unit fitted in aircraft outer skin opening. Window unit (1) to be fitted into opening (3) in aircraft outer skin (4) comprises frame (2, 34, 40) to receive package (6). Proposed unit comprises also frame (25, 36). Window frame (2, 34, 40) can be riveted to inner surface (50) of outer skin (4). Also one fastener (14) to secure pressure frame (25, 36) is riveted to window frame (2, 34, 40) and outer skin (4). Window frame (2, 34, 40) has L-like or rectangular cross section.

EFFECT: simplified design, increased glasing area.

13 cl, 5 dwg

FIELD: machine building.

SUBSTANCE: element of structure has aperture (4) in wall (3) of cavity where through moisture present in cavity can be drained into environment. A draining facility for pressurising aperture (4) in wall (3) of the cavity is provided in the aperture. The draining facility consists of hollow cylinder (6) connecting cavity of the element of the structure with environment; the element is equipped with external thread (7) having external diametre (G) and head (8) on one end with diametre (Ko). External diametre (G) of thread is less, than diametre (Ko) of the head.

EFFECT: reduced expenditures for compound parts and assembly equipment.

8 cl, 4 dwg

FIELD: transport.

SUBSTANCE: invention refers to aircraft industry, namely to protection device for aircraft door hinge. Aircraft hinge protection device is equipped with flexible covering (4) which can be tensed when the hinge is in closed position along the external hinge outline. Such construction provides closing of the openings between the first and the second elements (1, 2) of the hinge.

EFFECT: aerodynamic characteristics improvement.

11 cl, 12 dwg

Reinforced door // 2382719

FIELD: transportation.

SUBSTANCE: reinforced door intended to be used on board of aircraft contains internal structure with vertical pillars and connecting parts located horizontally. In this design internal structure is reinforced using net of belts connecting vertical pillars and connecting parts located horizontally.

EFFECT: flight safety improvement.

7 cl, 16 dwg

FIELD: information technologies.

SUBSTANCE: element of glasing comprises electronic device, intended for recording of information, which may be read at the distance with the help of information reading instrument. Electronic device comprises storage device, facility for reception and processing of information, at least one element of information detection and transfer, related to this element of glasing or to device, into which this element of glasing is inserted. Besides facility for reception and processing of information is connected to storage device so that to provide for transmission and recording of information into storage device in response to signals received at least from one element of detection. Method for registration of information and reading of information consists in collection of information with the help of detection element, transmission of this information into electronic facility for data reception and processing, registration of processed information with the help of electronic facilities in storage device, reading of information contained in storage device, with the help of contactless instrument for information reading.

EFFECT: invention provides for possibility to obtain information on technical characteristics related to element of glasing, obtained both at the stage of manufacturing and in process of this element operation.

19 cl, 1 dwg

FIELD: aircraft engineering.

SUBSTANCE: proposed system comprises a pair of transparent sheets (8) held apart and facing each other with the help of intermediate transparent layer (14) and resistive coating (20) arranged between aforesaid sheets. Proposed system incorporates also inverter (22) to feed AC square- or quasi-square signal (34) to aforesaid resistive coating (20). Said signal received, coating (20) emits heat to reduce or rule out water or ice accumulation of aircraft windscreen.

EFFECT: reduced weight and higher efficiency.

22 cl, 5 dwg

FIELD: shipbuilding.

SUBSTANCE: invention is related to aircraft window frames. Window frame comprises external flange, internal flange, vertical flange, transparent part of frame and fixing devices. Vertical flange is installed between external and internal flanges. External flange is intended to form connection of window frame to bearing structure of aircraft. Transparent part of window rests on internal flange or vertical flange and is retained with fixing devices, which are attached to window frame. Fixing devices are arranged at the same distance one from each other and are distributed along external flange. Fixing device comprises support plate, having surface, hold-down fixator and pin. Pin is arranged perpendicularly relative to surface of according support plate. Hold-down fixator is arranged as suitable in shape for engagement to pin. External surface of pin is equipped with geared structure. Hold-down fixator has elastic hook-like snapping elements. Geared structures of pin surfaces are intended for engagement to elastic hook-like snapping elements of according hold-down fixators.

EFFECT: invention provides for reduction of frame weight.

6 cl, 8 dwg

FIELD: mechanics.

SUBSTANCE: emergency exit opening system consists of a door and pneumatic cylinder connected with the door by means of a rod. The pneumatic cylinder is connected with one source of compressed air, as minimum, by pipeline through release mechanism. The release mechanism controls include a handle. The release mechanism contains one rocking mechanism coupled with the said handle and, as minimum, one pipeline cock, connected with the main rocking mechanism and retained in closed position of spring element. As minimum, there is one pyro-pusher in the release mechanism. The said pyro-pusher is installed so that rocking mechanism position can be changed when pusher is actuated. In addition, release mechanism controls include control element installed in the cabin and connected with the said pyro-pusher. Pneumatic cylinder is provided with rod fixing element in maximum door opening position.

EFFECT: increased reliability of door opening during flight opposite to flow and further door fixing.

8 cl, 2 dwg

FIELD: aeronautical engineering; canopy opening-closing mechanisms.

SUBSTANCE: proposed system includes swing section 2, fasteners for securing it on fuselage, suspension 3 and fasteners for securing the swing section to suspension. Suspension body has at least one first member secured on front wall of suspension and connected with second member 24 which is articulated with at least one rod 30 connected in its turn with fuselage by means of axle 18 and second member 24 connected with drive. Secured on suspension are first guides 16 limiting motion of rods, second guides 31 with horizontal and vertical sections for motion on fuselage rollers; at the rear there are brackets with slots for axle 18 and projections 13 engageable with limiters 19 mounted on fuselage. Drive is made in form of pneumatic hydraulic cylinder 9 with breaking rod. One motion of pneumatic hydraulic cylinder makes it possible to open, close or jettison swing section of canopy.

EFFECT: ease in use and high speed of response.

11 cl,5 dwg

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