• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    This invention relates to the field of aviation technology. The aim of the invention is to improve the performance by delaying the formation of intense shock waves and the separation of the boundary layer with increasing relative Mach numbers. The profile includes the upper 1 and lower 2 contours, convex from the leading edge 3 of the profile on most of its chord, concave near the trailing edge 4 and asymmetrically located relative to the chord of the profile connecting the front and rear edges. The coordinates of the points of the upper and lower contours in relation to the chord are given by a table for a profile having a maximum relative thickness of 3.5%. 2 hp ff, 7 ill.
    公开号:SU1540653A3
    申请号:SU833665258
    申请日:1983-11-17
    公开日:1990-01-30
    发明作者:Тибер Жан-Жак；Буске Жан-Марк
    申请人:Оффис Насьональ Дъэтюд Э Де Решерш Аэроспасьаль (Фирма)；
    IPC主号:
    专利说明:

    Faye. /

    s
    The invention relates to aeronautical engineering and relates to the aerodynamic profile of an aircraft propeller blade blade. The invention aims to improve performance by delaying the formation of intense shock waves and tearing off the boundary layer with increasing relative Mach numbers.
    Fig, 1 shows the aerodynamic profile of the propeller blade; Fig. 2 shows a symmetrical profile in which the law of variation in the thickness of the profile according to the invention is implemented; fig.Z - node I in figure 2; Figure 4 shows the profile change curvature curve; Fig. 5 illustrates the pressure distribution along the surface of the symmetric profile at, 88 and the NACA profile 16; figure 6 - the same, with, 92; Fig. 7 illustrates the change in the drag coefficient as a function of the number M for the profile according to the invention and the profile NACA 16304.
    The aerodynamic profile includes the upper 1 and lower 2 contours, convex from the leading edge 3 on the greater part of its chord and concave near the trailing edge 4.
    The aerodynamic profile has four successively located zones I, II, III, IV, in which a certain law of thickness variation from a profile is observed, and three zones 1a, Ha, Ilia, in which a certain law of variation of curvature f of the middle profile line is observed.
    The x-axis of the rectangular coordinate system coincides with the chord of the profile and is positively directed from the leading edge 3 to the trailing edge 4, with the beginning coinciding with the leading edge 3
    The OY axis is positively directed from the lower to the upper contour.
    For a symmetric profile (figure 2), the front edge may have the shape of an arc (figure 3), while the law of thickness variation C profile between the front edge 3 and the section of the profile with the greatest thickness CMOKt (end of zone II) can be expressed by the formula
    U4x6 + U2xs + U3x4
    U4xs

    7, (Oh, the constant coefficients are related to the chord b.
    In the case when zone II goes up to 32% of the chord and when the maximum relative thickness C / B is 0.035 (3.5%), the coefficients of formula (1) have the following values:
    U, - -44,937
    Ut - 49.052
    U3 -19,232
    U. - 3.0433
    U5-0.1028
    U6 - -0.0720
    U7-0.0649.
    Zones III and IV can be represented by formula (2) between the thickest section and the trailing edge:
    0
    five
    0
    Uex7
    + U
    权利要求:
    Claims (3)
    [1]
    I
    and
    + and
    ieh + ts0h and
    + and ..
    x +
    (five
    U, 4x + U
    five
    (2).
    The constant coefficients U8 - U, j are chosen depending on the coefficient. (1) so as to ensure the continuity of the profile at point 5 of maximum thickness. In this case, when the relative thickness is 035 (3.5%), and the thickest section is located at 32% of the chord, they can have the following values:
    0
    35 40
    45
    50
    U - U, Uy U "
    five
    8.4058 -34.3764 58.4983
    U, 4 -53.5964
    Utt 28,4971
    U, 3 -8.8134 0.1472 -0.0852.
    The beneficial effect of the above-defined thickness variation is advantageously complemented by the law of curvature variation, which allows excellent results to be obtained at elevated values of the carrier force for quantities (M numbers about 0.6, which corresponds to the take-off and climb modes.
    The law of curvature can be represented, like the law of thickness, by a curve in the system of axes, where the axis OX coincides with the chord b, and the axis of ordinates OY (figure 4) is directed from the lower surface of the blade to its upper surface. The law of curvature can be decomposed into three zones represented by functions, for example, polynomial.
    The first zone 1D between the front crown and the point of maximum curvature Pb can be approximately determined by the type relation
    71
    Invention Formula
    The aerodynamic profile of the propeller blade of an aircraft, including upper and lower contours, convex from the leading edge of the profile on most of its chord and concave near the trailing edge and located asymmetrically relative to the chord of the profile connecting the front and rear edges, which differs from in order to improve performance by delaying the formation of intense shock waves and the separation of the boundary layer with increasing relative Mach numbers, the coordinates of the points of the upper and lower contact trench defined ratios
    Yu X / ъ-Ј "X / Ь and YH" / b-f2 (x / b),
    where b is the chord of the profile;
    x / b is the ratio of the coordinates of the points of the upper and lower contours along the X axis, which coincides with the chord, to the chord;
    TL / L is the ratio of the coordinates of the points of the upper contour along the Y axis, perpendicular to the X axis, the origin of which is located at the leading edge of the chord;
    H1 / b - the ratio of the coordinates of the points of the lower contour along the Y axis to the chord,
    moreover, the values of x / b, Ґ6epx / b, Ґmiž / b
    are shown in the table
    [2]
    2. Profile pop. 1, which differs by the fact that the maximum relative thickness of the profile is from 2 to 6% of the chord.
    [3]
    3. The profile according to claim 1, characterized in that the maximum curvature of the centerline of the profile is located on 35% of the profile chord.
    ; / 0.08
    Фа2.1
    maximum curvature is 35% of the chord, i.e. Pb 0.35b (35% of the chord), and the maximum curvature value взMax taken approximately equal to 0.01361), the coefficients can have approximately the following values:
    M (0.1589 Mt -0.2151
    M5
    M4
    m, "
    m,
    M7
    ma mg
    V
    V
    m "
    0.0730
    0,00252 -0,01869
    0.03069 -0.03876
    0.02676
    0.0874 0.00455 0.0694 -0.1411
    Combining the law of change in the curvature of the midline and the law of change in the thickness of the profile makes it possible to obtain an asymmetric profile shown in Fig. 1.
    The use of functions (1) - (5) and the coefficients U and M indicated above determines the coordinates of the points of the upper
    thirty
    35
    Obtaining profiles, the relative thickness of which is in the range from 2 to 6%, is performed by simply multiplying the ordinates listed in the table by the ratio of the relative thickness of the desired profile to the relative thickness of the 3.5% profile, the coordinates of which are listed in the table. I
    The pattern of pressure distribution along the surface of the symmetric profile (curve 5, Fig. 5) shows an increased value of the pressure coefficient Cp to approximately 20% of the chord and keeping it at a level exceeding the value of this coefficient for the profile type NACA16 (curve 6) with the number 88. Pressure distribution with increasing number, 92 is shown.
    and the lower contours, given by the ratio of -40 to This allows to obtain the following:
    it is intense and located near the front end of the shock waves, as a result of which the separation of the boundary layer is delayed and the level of drag is reduced at large values of M "
    Obtaining profiles, the relative thickness of which is in the range from 2 to 6%, is performed by simply multiplying the ordinates listed in the table by the ratio of the relative thickness of the desired profile to the relative thickness of the 3.5% profile, the coordinates of which are listed in the table. I
    The pattern of pressure distribution along the surface of the symmetric profile (curve 5, Fig. 5) shows an increased value of the pressure coefficient Cp to approximately 20% of the chord and keeping it at a level exceeding the value of this coefficient for the profile type NACA16 (curve 6) with the number 88. Pressure distribution with increasing number, 92 is shown.
    This allows you to get
    it is intense and located near the front end of the shock waves, as a result of which the separation of the boundary layer is delayed and the level of drag is reduced at large values of M "
    Comparative tests of the profile shown in Fig. 1 with a maximum relative thickness of 3.5% and a profile of the type NACA 16304 with a maximum relative thickness of 4%, carried out in a wind tunnel (Fig. 7), show the UTO increase in drag depending on the number M for the profile according to the invention (curve 7) is significantly delayed compared to the profile of JASA 16304 (curve 8). The gain is about 5%.
    Y
    league
    FIG. 3
    1.0 / AT
    0.5
    Y ", 5
    $
    0.5
    ABOUT
    Or
    0.8
    0.85
    1.0
    Qf
    N
    x / 8
    FIG. 6
    0.9 fm.7
    0.95
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    同族专利:
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    FR2536365B1|1985-04-26|
    FR2536365A1|1984-05-25|
    DE3372195D1|1987-07-30|
    JPS59118597A|1984-07-09|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    WO1994008846A1|1992-10-16|1994-04-28|Aktsionernoe Obschestvo 'aviatika'|Propeller|US3343512A|1966-05-20|1967-09-26|Francis R Rasmussen|Hydrofoil with unsymmetrical nose profile|
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    US4142837A|1977-11-11|1979-03-06|United Technologies Corporation|Helicopter blade|
    GB2016397B|1978-02-02|1982-03-24|Aerospatiale|Aerofoil|
    US4459083A|1979-03-06|1984-07-10|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Shapes for rotating airfoils|
    FR2463054B1|1979-08-10|1983-08-12|Aerospatiale|
    FR2490586B1|1980-09-24|1982-10-01|Aerospatiale|
    US4412664A|1982-06-25|1983-11-01|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Family of airfoil shapes for rotating blades|USRE33589E|1986-09-03|1991-05-14|United Technologies Corporation|Helicopter blade airfoil|
    US4744728A|1986-09-03|1988-05-17|United Technologies Corporation|Helicopter blade airfoil|
    US4834617A|1987-09-03|1989-05-30|United Technologies Corporation|Airfoiled blade|
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    GB2220712B|1988-07-13|1992-12-09|Rolls Royce Plc|Open rotor blading|
    US5024396A|1988-07-19|1991-06-18|Principia Recherche Developpement Sa|Air or submarine engine with improved contour|
    US4941803A|1989-02-01|1990-07-17|United Technologies Corporation|Airfoiled blade|
    JP2633413B2|1991-06-03|1997-07-23|富士重工業株式会社|Rotor blades of rotary wing aircraft|
    FR2689852B1|1992-04-09|1994-06-17|Eurocopter France|BLADE FOR AIRCRAFT TURNING WING, AT THE ARROW END.|
    US5417548A|1994-01-14|1995-05-23|Midwest Research Institute|Root region airfoil for wind turbine|
    US6378802B1|1998-05-04|2002-04-30|Manuel Munoz Saiz|Enhance aerodynamic profile|
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    法律状态:
    优先权:
    申请号 | 申请日 | 专利标题
    FR8219337A|FR2536365B1|1982-11-18|1982-11-18|
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