KR820002096B1 - Helicopter blade - Google Patents

Abstract

The helicopter blade(10) has a central portion(18) extending between the base(16), fitted to the rotor hub(12), & the tip(20). There can be stabilizers at the trailing edge to eliminate the pitching moment. The maximum negative pressure acting on the top surface can be exerted on a front portion of the chord equal to 10% of the chord length(28), and the blade tip can be offset to the rear in relation to the central part. This offset to the rear introduces an effective sweep-back of the tip, which delays compressibility drag rise on the advancing blade.

Description

Helicopter blade

1 is a plan view of a helicopter blade of the present invention,

2 is a cross-sectional view of the helicopter blade of the present invention,

3 is a graph of a local Mach number of Ikhyun comparing blades of the present invention with respect to SC1095 blades,

4 is a graph of the maximum lift coefficient for Mach number of free air flow comparing the blade of the present invention with respect to the SC1095 blade,

5a and 5b are graphs of drag versus Mach number when lift coefficients are 0.1 and 0.9, respectively.

6 is a graph of the maximum positive ratio for Mach numbers,

7 is a graph of lift ratio versus lift coefficient.

The present invention relates to a helicopter blade, and more particularly to a helicopter blade having a cross section that can improve the stall of the blade.

Helicopter blades, hereinafter referred to as blade SC1095, are described in US Pat. No. 3,728,045, with the most recent known prior art advantages over prior art techniques.

The first object of the present invention is to reduce the magnitude of the peak pressure on the upper side of the airfoil type blade and distribute the peak pressure to the blade blades of the blade so as to have a cross section to reduce the flow separation and the increased drag thereby. It is to provide an improved helicopter blade that has advantages over blades.

It is another object of the present invention to provide a blade having a high drag ratio (D) and having no excessive shock wave generation, which reduces the air flow rate on the upper surface of the blade from the blade trailing edge to below the speed of sound. This is accomplished by selectively shaping the upper curvature of the blades to compress the air flow that gradually produces the peak pressure.

When described in more detail by the accompanying drawings illustrating the present invention.

The blade of the present invention entails all the advantages of the SC1095 blade of US Pat. No. 3,728,045, which has advantages over the prior art, while increasing maneuverability of the aircraft, reducing the load on the steering rods and reducing vibration. Therefore, in the present specification, the blade of the present invention and the SC1095 blade were compared to show the improvement of the blade of the present invention.

1 and 2 show the invention helicopter blade 10 connected to the hub 12 of the rotatable shaft 14, the blade 10 being connected to the hub 12 in a general manner. The blade 10 is composed of the root portion 16, the central action portion 18 and the tip (20), the leading edge 22 and the trailing edge 24 extends backwards, the blade length in the figure In (26), the magnitude of the chord is indicated by reference numeral 28.

The SC1095 blade of US Pat. No. 3,728,045 poses a threat of stall stalling under certain flight conditions. However, while the blade of the present invention retains the other advantages of the SC1095 blade, this threat is raised due to the stall problem. The present invention reduces the peak pressure of the upper side of the airfoil-type blade and distributes the peak pressure to the upper blade edge of the blade to increase the maximum lift, eliminate the separation of air flow, and selectively reduce the drag. By modifying the foil cross section, stall problems can be eliminated.

Figure 3 illustrates the advantages of the blades of the present invention, the blades of the present invention being labeled R8 blades for the SC1095 blade. 3 is a graph showing the constant lift coefficient of the local Mach number according to the blade size of the blade. Here, it can be seen that the negative pressure peak on the upper side of the blade of the SC1095 blade is larger than the blade of the present invention, and in the SC1095 blade, it actually causes a shock wave. It can be seen that the pressure gradient between A and point B is very large. This shock wave reduces the flow velocity through the upper side of the blade to a point below the sonic speed of the blade rear face 24. The negative pressure pattern of the SC1095 blade shown in FIG. 3 is important for generating lift of the blade. However, as will be described later, it is possible to enjoy the peak pressure by the selective deformation of the R8 blade of the present invention, and thus to reduce the magnitude of the shock wave according to the magnitude of the pressure hardness, so that the peak pressure can be dispersed on the blade vane phenomenon of the blade. This avoids the formation of supercritical shockwaves on the top surface of the SC1095 blade. In addition, since the negative pressure at the lower extremity of the R8 blade of the present invention is greater than the negative pressure of the SC1095 blade, it is possible to eliminate or greatly reduce the flow separation and thereby reduce the drag of the R8 blade structure of the present invention. L / D) can be improved beyond the SC1095 blade. As can be seen from the graph of Fig. 3, it is possible to change the local Mach number to the pressure coefficient C _p and also to change the display of the coefficient, so it is understood that the pressure shown on the upper side of the blade actually represents negative pressure. It is important.

As shown in FIG. 3, the maximum negative pressure of the blade of the present invention occurs at 10% front of the blade.

The superiority of the R8 blade of the present invention over the SC1095 blade is also illustrated in FIG. 4 in which the free airflow Mach number for the maximum lift coefficient C _Lmax is shown. As shown in FIG. 4, the R8 blade of the present invention actually exhibits a large maximum lift coefficient at a lower Mach number than the SC1095 blade.

Figures 5a and 5b compare the different performances between the two blades, in Figure 5a the drag of the SC1095 blade is constant over a given Mach number range and the drag of the R8 blade of the present invention is the same as the drag of the SC1095 blade. One indicates higher mach numbers. Figure 5a shows the action of the blade tip as shown at a very low lift coefficient with a lift coefficient C _L of 0.1. However, in the blade of the present invention, the blade tip uses the same cross section of the SC1095 blade tip, but the blade tip acts as a high Mach number. If the pitching moment is a problem, the R8 blade of the present invention can be used by removing the general trailing edge step 30.

In FIG. 5B, the lift and Mach number relationship is shown when the lift coefficient C _L is 0.9, which is very high, and shows the action of the blade center operating portion between the blade tip and the blade root. Here the blade airflow is approximately. The Mach number range is between 3-.5 and the drag on both blades is about the same.

6 shows the maximum drag ratio (L / D) max for the Mach number, which shows the improved performance of the R8 blade of the present invention over the SC1095 blade in the range between Mach number .3-.5. The R8 blades showed an increase in the maximum lift coefficient of .3 but only a slight weakness at high drag and pitching moments.

In general, the blade of the present invention is tapered in the thickness of the inner side and the thickness of the outer side, so that a thick blade tip is not required and the thin tip acts to reduce drag in the high Mach number.

Experience has shown that a blade lip is needed to perform a small lift function, but the blade's central action actually produces all the lift required. Based on this fact, the SC1095 blade generates less drag, so the present invention R8 blade shown in FIG. 1 is practically used between 40% and 80% of its blade length, and the cross section of the SC1095 blade has a blade tip and a blade root. Used in

7 shows a comparison of the different performances of both blades, showing the lift ratio with respect to the lift coefficient C _L corrected for the pitching moment, thereby demonstrating the structural superiority of the R8 blade of the present invention.

4 and 6 show the maximum lift coefficient performance and drag ratio (L / D) performance of the R8 blade of the present invention better than the SC1095 blade, which has blade stall and accompanying problems and reduced maneuverability and increased performance in the SC1095 blade. It is very important in view of the problem of increased control load and increased vibration, and all these problems can be reduced or eliminated by the structure of the blade of the present invention.

The R8 blade of the present invention is illustrated by the coordinates and equations as shown in FIG. 2. The R8 blade of the present invention has a front camber having a thickness of about 9.4% manifestation size and a maximum camber of about 1% manifestation size. Located at 30% shear, the pitching moment coefficient of the blade is within the range of ± .03, and the low peak pitching moment can be obtained without loss of high maximum lift coefficient and drag dissipation at all Mach numbers below .75 before moment divergence.

In general, the airfoil shape is determined by defining the lower airfoil surface at the continuous surface end along the upper airfoil surface and the chord and the radius of the leading edge. This is detailed, for example, in Abbott and Wen Dohnhofer's "Theory of Wing Selections", published by Dover Publishing Co., New York in 1959, and the standard way of determining airfoils is described in detail in page 412 of the book. It is.

The following table determines the inventive airfoils for blade thickness.

Where X is the position of the chord, C is the chord, Y _u is the coordinate or position of the upper airfoil surface from the chord at position X, or Y _L is the coordinate or position of the lower airfoil from the line of the chord at position X, t is the maximum blade thickness.

It is possible to set the positions Y _u and Y _L of the upper and lower air foil surfaces at any wing position X for the blades of the selected thickness and the selected ship size.

The following example illustrates the use of the table above to determine the position of Y _u and Y _L for a leading distance of position 0.0125 along the beginning of the leading edge at the leading edge, with the blade having a maximum thickness of 2 inches and the leading edge of 20 inches.

Determining the Y _u is given enemy (積) of 0.03726 inchi by multiplying the maximum thickness, i.e., two inches of the blade in the Y _u / t term 0.1865 0.0125 for the chord locations. This less indicates the position of the upper airfoil, that is, Y _n . Therefore, the position Y _u of the upper airfoil at the blade position 0.0125 is 0.03726 inches on the ship's ship.

Using this calculation method, the distance Y _L can be found at the star position 0.0125, and since the items in the Y _L / t column represent negative numbers, the figure is located below the line of the star at the port position 0.0125. According to this calculation method, Yu and Y _L of all star positions are set in the X / C column.

It is also possible to determine the total radius of the upper airfoil surface P _{u and} the total radius of the lower airfoil surface P _L. Therefore, this calculation method is calculated by multiplying in two stages. In the first stage, the ratio of the maximum thickness of the blade to the leading edge [t / c = 2 ^N / 20 ^N = 0.1] is multiplied by 1.108.

That is, (0.1) 2 is multiplied by 1.108 to obtain 0.01108. This first step represents the total air radius P _u divided by the extrudate C. The second stage is calculated by multiplying the enemy's first stage by the size of the chord, the pole 0.01108 by 20 inches. The second stage's enemy is 0.2216, which yields the leading edge radius of the upper airfoil surface at the size of the chord. The lower leading radius P _L is calculated in the same way.

As described above, the numerical values listed in Y _u / t and Y _L / t can be regarded as the size of the chord. If you determine the coordinates for blades with different starch numbers, you need to multiply the number of items in each of these two columns with the starch size.

By using the table given above, the coordinates of the cross section of the airfoil of the blade can be determined and the above-listed advantages are obtained by this airfoil when the numerical value of the item varies in the range of ± 3%.

Since the blade R8 of the present invention is not listed in the standard NACA standard, the following table can be used to set the upper and lower airfoil positions Y _n and YL at the position X / C of the blade string.

Here, the symbol X is the position of the blade chord, C is the size of the chord, Y _u is the position of the upper airfoil surface, Y _L is the position of the lower hairfoil surface.

This second table, which differs from the first table, is the coordinate for the R8 blades with a thickness to tip ratio (t / c) of 9.4%. The first table given is more fluid, and by using the calculation method described in this regard, the upper and lower airfoil positions can be determined as the ratio of the thickness to the tip.

The airfoil cross section can also be described by the following equation.

Upper side

Lower side

Where X is the position of the chord, C is the size of the chord, Y _u is the distance of the upper airfoil at the tip X of the tip, t is the maximum thickness of the blade, and Y _L is the surface of the lower airfoil at the tip X from the chord. Indicates the distance of.

Similarly to the table given above, the portion of the airfoil that benefits from the present invention is obtained by finding all necessary coordinates Y _u and Y _L for each extreme position X within a range of ± 3% of these Y _u and Y _L values. It can be calculated by introducing into.

The present invention is not limited to the exact details, and it is believed that those skilled in the art will be able to suggest various modifications.

Claims (1)

A helicopter blade having a thickness, span, and chord, comprising: a root portion adapted to be connected to a rotahub, a tip portion forming a blade portion furthest from the root portion, and the root portion; Helicopter blades, characterized in that consisting of a central portion in the form of the form of the airfoil of the cross section is defined by the following coordinates attached and connected between the tip and the blade.

Where X is the position along the blade chord, C is the dimension of the blade chord, Y _u is the position of the upper airfoil surface, Y _L represents the position of the lower airfoil surface.

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