JPH02155721A - Manufacture of blade - Google Patents

Abstract

PURPOSE: To obtain an efficient propeller blade constituted from a unidirectional layer of a fiber reinforced resin-bonded structural composite material by heating the blade during a proper curing cycle for a matrix of low modulus of elasticity and bonding a blade beam, a foaming, filler, the first and second shells together. CONSTITUTION: A projected front fringe 414 and a recessed rear fringe 416 are formed by crossing a front surface 409 and a rear surface 407 to form retreat blade having a radial axis 422. These surfaces form a shell composite, a metal blade beam 418 is arranged between the surface 409 and the rear surface 407 to be bonded to the shells with the use of a structural adhesive. A foaming packing space 424 is formed near the front fringe of the blade beam 418, and a rear side foaming packing space 426 is formed near the rear fringe. Bonding members 420 are inserted into the shell and the blade beam to bind to prevent the separation of the blade. The second bonding members are inserted into the vicinity of the blade beam through the shell to fasten the shell to the blade beam. The front side and rear side blades are constituted from a unidirectional fiber layer, and the fibers are put in a matrix material having ductility, low strength, and a low modulus of elasticity.

Description

DETAILED DESCRIPTION OF THE INVENTION Field of the Invention This invention relates to aircraft propulsion systems and, more particularly, to a method of making very thin, high sweep, wide chord propeller blades primarily from composite materials.

BACKGROUND OF THE INVENTION Basic aircraft propellers are conventionally made of structural materials, but these materials have been shown to increase only one of the key parameters of structural efficiency, such as strength or stiffness. , have not reached their expected abilities. Such materials include steel, wood, aluminum, titanium, etc. Because of these drawbacks, the aircraft industry has developed a strong interest in high-grade fiber reinforced composites. If used correctly, these materials can reduce the weight of the blade by as much as 50% while increasing its strength.

The aircraft blades considered in this invention are those in a counter-rotating propeller system having a front propeller with 5 to 15 blades and a counter-rotating rear propeller with 5 to 15 blades. These blades have a large degree of retraction,
The strings are wide and very thin. Propeller blade airfoils operate at transonic and supersonic speeds.

Traditional propeller blade designs have been adequate for subsonic flight. However, when these blades were used for high subsonic flight, a number of structural problems arose that reduced performance. Structural problems that occur with blades operating at very high speeds are caused by the forces and stresses acting on the blades. The forces acting on the blade during flight are thrust, centrifugal force, and torsional force. First, the thrust induces bending stresses in the blades. Second, centrifugal force stretches the vanes radially. Finally, torsional forces force the vane to twist about the vane's radial axis. The ideal feather is
In order to improve fuel consumption efficiency, these forces are suppressed while minimizing weight.

The solution to the wing problem is fiber reinforcement1. This was the development of a resin-bonded structural composite material. These materials have created new design possibilities for propellers. There are three main advantages to using fiber-reinforced composites. First, complex airfoil shapes can be molded. Second, composite old materials save weight. Third, the dynamic frequency response of the vane element can be tailored to its operating parameters. The present invention overcomes the problems and deficiencies of conventional blades by providing a propeller blade constructed from a composite material with strength and airfoil shape to provide an efficient blade for counter-rotating propeller systems.

SUMMARY OF THE INVENTION It is an object of the present invention to provide an efficient propeller blade constructed from a unidirectional layer of fiber-reinforced resin-bonded structural composite material.

Another object of this invention is to provide a counter-rotating propeller blade for an aircraft that overcomes the forces and stresses to which the blade is subjected at high speeds.

Another object of the invention is to provide a counter-rotating propeller blade for an aircraft which improves the efficiency of the counter-rotating propeller system.

Another object of the invention is to reduce vibrations in the device and
It is an object of the present invention to provide a blade with radial and chordal balance in order to reduce the bending load at the root of the blade.

Generally speaking, aircraft propeller blades are constructed from a plurality of angular stacked composite laminates. These composite laminates form first and second shells with intersecting surfaces at the leading edge, trailing edge, root section, and tip. The leading and trailing edges are swept back to reduce the noise generated at the tip and to reduce aerodynamic losses due to air compression effects. A metal wing spar is interposed between the first and second shells and bonded to the surface to strengthen the surface. To reduce the weight of the vane, a void is provided between the faces. Adjustable position counterweights in the spar on the forward and aft sides of the spar provide radial and chordal balance of the blades.

The blades are constructed by providing a pattern of blades such that, for each side of the shell, a region of constant thickness is obtained in the cross section of the blade. The composite material is molded to the contours of this pattern. A first shell is formed by overlapping multiple composite layers of molded material. The composite material includes unidirectional fibers encased in a low modulus matrix. The fibers in each layer are aligned in varying directions to provide strength and stiffness to the blade. A second shell is similarly formed. Aligning the wing spar between the first and second shells. The foam filler is then placed in the predetermined location to form a void between the shells. Apply adhesive to the wing spar, inner surface of the shell and foam filler.

The adhesive, spars, shell and foam filler are then bonded together by curing in an autoclave with a curing cycle appropriate for the low modulus matrix.

A bonding member is then inserted into the composite vane to tighten the shell, spars and foam filler. The blades are then radially and chordally balanced to minimize vibrations in the device and bending loads at the root of the blades.

Other objects of the invention and advantages thereof will become apparent from the following description of the drawings.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS FIG. 1 shows the counter-rotating propeller system of the present invention. - For boat fishing, the first propeller 100 includes a front blade 106 rotating in the direction indicated at 102 and the rear propeller 104 includes a rear blade 108 rotating in the opposite direction 105. All blades of the front propeller 100 are the same, and the blades of the rear propeller 1
All vanes of 04 are also the same. The front blade and the rear blade have different dimensions. Typical leading and trailing vane constructions are described in pending US Patent Application Serial No. 157,179, filed February 16, 1988. This is a now abandoned U.S. Patent Application Serial No. 932.4 filed on November 19, 1986.
This is a continuation application No. 27.

Referring to FIG. 2, the forward vane 106 has an airfoil portion 310 that includes a tip 312 and a root end portion 302. As shown in FIG. The airfoil portion 310 is the wing tip 3
12 and the root portion 302 between the front surface 309 and the rear surface 30
7 (this is hidden in the figure) and is composed of a plurality of angular stacked composite laminates with continuous fibers embedded in a matrix material. The front edge 314 has a convex shape where the front surface 309 and the rear surface 307 intersect.
and a concave shaped trailing edge 316 to form a swept vane having a radial axis 322 as shown in FIG. Front 3
09 is convex, but the rear surface is concave. Two shells form these surfaces, with a metal spar 318 interposed between the front surface 309 and the rear surface 307 and bonded to the shells to connect the surfaces with the root portion 302 . A forward foam fill cavity 324 is provided near the leading edge of the wing spar 318 . An aft foam fill cavity 326 is provided near the trailing edge of the wing spar 318 .

A leading edge sheath 311 is attached to the leading edge of vane 106 to protect it from erosion. The plurality of coupling members 320 are surfaces,
Passed through the shell and wing spar to restrain the blades from separating under strong loads. A number of coupling members are inserted adjacent the leading and trailing edges of the spar to clamp the surfaces to the spar.

FIG. 3 shows the structure of the rear wing. Generally speaking, the rear vanes 108 are similar in shape to the front vanes 106. The trailing blade 108 has an airfoil portion 410 that includes a tip 412 and a root portion 402 . The airfoil section 410 has a front surface 409 and a rear surface 407 (hidden in the drawing) between the wing tip 412 and the root section 402, which form an angular structure composed of continuous fibers embedded in a matrix material. It is composed of multiple composite laminates stacked one on top of the other. Continuous fibers of the composite laminate extend throughout the Aerofoil.

A front edge 41 having a convex shape where the front surface 409 and the rear surface 407 intersect
4 and a concave trailing edge 416, which forms a swept vane with a radial axis 422 as shown in FIG. The anterior surface 409 is convex, while the posterior surface 407 is concave. These surfaces form a shell composite, with a metal wing spar 418 interposed between the front surface 409 and the rear surface 407, and a structural adhesive material such as AF3109-2 manufactured by 3M Corporation. is used to connect to the shell. Note that adhesive is used to join the front vanes as well as the rear vanes. A foam fill cavity 424 is provided near the leading edge of the wing spar 418.

An aft foam fill cavity 426 is provided near the trailing edge of the wing spar 418 . A leading edge sheath is bonded to leading edge 411 for protection. A plurality of coupling members 420 are inserted into the shell and spar to restrain the blades from separation. A second plurality of coupling members are inserted through the shell and proximate the spar to tighten the shell to the spar.

The front and rear vanes each consist of a layer of unidirectional fibers.

The fibers are unidirectional and housed side-by-side in parallel in a ductile, low strength, low modulus matrix material.

This matrix material localizes the effects of a single fiber break by transmitting the load from fiber to fiber through shear, redistributing the load near the ends of a broken fiber to adjacent fibers. Ru. A typical fiber used in practicing this invention is a composite of 80% graphite and 20% S-glass. However, combinations of various fibers can be used, including, for example) [evlar(R)] fibers, boron fibers, glass fibers, and the like.

This invention will be specifically explained below regarding the front blade.
At the outset of this description, it should be noted that the apparatus and method of the present invention also apply to the trailing vanes. Furthermore,
More than one type of fiber may be used in each layer or combination of layers.

+80@, +3 from the radial axis 322 of the fibers in each layer
Layer the stack by aligning in an alternating pattern of 5''-10'' and +35@. Two successive layers may be stacked at the same angle. Note, however, that when the angle of a layer is changed, the order described above is followed. As will be apparent to those skilled in the art, this angular order can be varied to create a composite laminate with desired strength properties in various directions of the vane.

Such a blade configuration provides an aeroelastically stable blade with well-tuned vibration modes.

Referring to FIG. 4, the central wing spar 31 of each blade
8 is made of high strength metal such as titanium.

A metal wing spar 318 is interposed between the front and rear surfaces;
A suitable adhesive is applied to each and bonded to each side.

This adhesive provides extra bond strength to resist the blades from separating. The spar 318 strengthens the airfoil section and also transfers loads from the airfoil section to the root section, including the dovetail (not shown).

This dovetail, which secures the spar to the rotor hub, is well known.

One of the many functions of the spar 318 is to create a connection between the blades and the rotor hub. This wing spar is broken line 30
4,306, it has a hollow space created by a hollow undercut.

The undercut cavity accommodates a weight member that statically balances the vane relative to the radial axis 322 and the chordal axis. In the undercut cavity 304 of the forward spar is a counterweight 340 constrained to move within a groove 341. The counterweight 340 is in the groove 341
Because of the angle it makes with the radial axis and the chord, it has chordal and radial translation components. Furthermore, the rear counterbalance A34
2 is in groove 346 and has chordal and radial movement components.

The front counterbalance [11! 340 and the rear counterweight 342 are located in the front undercut cavity 304 and the rear undercut cavity 306, respectively. A counterweight moves inside the groove to provide radial and chordal balance of the vanes. This movement can be effected in the groove by providing a thread on the weight. However, other methods of moving the weights can also be used, as is well known to those skilled in the art.

Additional weights may be used to further change the mass of the vanes. These weights substantially change the center of mass or geometric center of the vane.

When making the blade, patterns are created for both the concave side 501 and the convex side 503 of the blade, as shown in FIGS. 5 and 6. Unidirectional, side-by-side parallel fibers in a ductile, low-strength, low-modulus matrix material are laid in one layer of this pattern and cut to fit the shape of the pattern in that layer. The low modulus matrix material used in practicing this invention is an epoxy resin. 5 and 6 show the concave and convex shell patterns of the blade, respectively.

The contour represents, for each side, an area of constant thickness in the cross section of the blade. Substantially, such a contour corresponds to a pattern for one layer of fibrous material.

For example, line 500 represents the pattern for the innermost portion of the vane, and fi 502 represents the pattern for the outer layer of the vane. When making wings, unidirectional, side-by-side parallel fibers in a ductile, low-strength, low-modulus matrix material are placed in this pattern, cut to match the pattern, and the direction of the fibers is determined. so that they align in the planned direction. When practicing this invention, the fibers are +80 from the radial axis 322.
Align in an alternating pattern of "+35" -109 and +35". Construct both halves or both sides of the vane by placing each layer on top of each other. Epoxy-impregnated layers are very sticky, so each layer The two shells or halves, one for the convex side and one for the concave side, thus form a wing shape.

Referring to FIG. 7, a first shell is introduced into mold 510. This mold is preformed to match the shape of the shell. The wing spar 318 is shaped like 31 as is well known to those skilled in the art.
By tightening the alignment mechanism at 8, the radial axis 3
22.

A blowing agent 5 is placed in the form of a cavity adjacent to the leading edge of the wing spar 318.
12, a leading edge cavity is formed. A trailing edge cavity is formed by positioning the blowing agent 514 in the cavity adjacent the trailing edge of the wing spar 318 . wing spar 3
Apply extra glue to each side of the wing spar to the first
is firmly coupled to the shell 502 of. Second shell 503
is placed over and aligned over the first shell, wing spar and foam filler and encapsulated in a mold. The mold is then heated to a predetermined temperature for a predetermined length of time depending on the curing cycle for the epoxy resin used in its manufacture. Remove the wings from the mold and let them cool. Trim excess resin from the edges and base of the blade. Attach the leading edge protector using well-known vacuum backing. A coupling member is then inserted into the vane from one side to the other to provide extra strength to prevent the vane elements from separating when loaded. Of course, there are many other desirable methods of making composite structures. The manufacturing method for producing large-volume blades is carried out by automated methods. However, it goes without saying that this automated method corresponds closely to the method just described.

After the blade is made, its static balance is established with respect to the two axes of moment. To minimize rotational unbalance caused by misalignment of the geometric centers of rotating blades. As an example, two vanes are shown rotating about centerline 12 in FIG. A rotational imbalance occurs when the graphic center 22 of the vane 24 is not radially and axially aligned with the graphic center 25 of the vane 26. For example, radial distance! If 2g is greater than the radial distance 30, a radial imbalance appears. Furthermore, if the vanes are not aligned in the same axial position, an imbalance will occur.

The axial position is defined as the position on the axis (center line 12) where the vertical line passing through the center of the figure intersects the axis. Any unbalance will cause vibrations in the propeller unit, which will cause power loss and, in extreme cases, destruction of the propeller unit. Therefore, in order to reduce vibration, rotational unbalance (MW
+ -MW2). Here, MW+ is the mass and weight of the first blade in Figure 8, and MW2 is the second blade.
is the mass and weight of the blade.

In order to eliminate this problem, the radial and axial directions (
balance in the string direction). The first axis is the radial axis and the second axis is the chordal axis. Referring to FIG. 9, the blade is captured by the weighing device 10. This weighing device may be a simple balance weighing device having a weight 12 for zeroing the device and a scale 4 for measuring the weight of the blade. In a well-known manner, the dead or areal weight of the blade is first determined.

Next, determine the amount of moment ffl relative to the radial axis. The moment weight in the radial direction is W x + where W is the dish weight and Xl is the distance from the fulcrum to the figure center 32. This moment weight is indicated on the scale 14. Since the radial moment weight and surface weight are known, Xl is determined. Determine XI for multiple blades of the propeller. Here, I represents the first blade. Next, determine the scattering of chordal moment distances for the propeller blades.

Referring to FIG. 10, the same weighing device ffl0 determines the chordal moment. The blade is rotated 90@ so that the axial position of the figure center 32 can be determined. Where W is the surface weight and Yr is the axial position of the center of the figure relative to the first blade, the chordal moment weight is WY
Equal to T. Movement of the front and rear counterweights may change the axial position Yl. Select the planned Yl for multiple blades. Thereafter, each blade is corrected so that the axial position of the center of the figure becomes the same for each blade. To this end, the YT is adjusted by carefully positioning, adding or subtracting the front and rear counterweights of each vane. After modifying the blades so that the axial positions of the graphic centers are the same, the radial positions of the graphic centers are aligned. For this reason, a weight is added or deleted at the base while maintaining the axial position of the center of the figure. For example, add 1 ounce to the base, add 1 ounce to the front counterweight, and add 1 ounce to the back counterweight.
Assuming you add oz. By doing this, the radial position of the figure center is changed while maintaining the axial position of the figure center. In the present invention, it has been found that this two-axis static balancing method provides adequate dynamic balancing of the rotating blades.

The method for manufacturing a counter-rotating propeller device for an aircraft having very thin blades with a large chord width and a large degree of retreat has been described above. The blades are made using two shells constructed from composite laminated materials. A composite laminate includes a plurality of unidirectional fibers in a low modulus matrix. A wing spar is interposed between the shells. Add adhesive between the wing spar and shell to increase bond strength. A foam filler is placed in a predetermined location between the shells to form a cavity within the vane. Curing in an autoflave with an appropriate curing cycle for the matrix bonds the shell, spars, foam filler, and adhesive.

Connecting members, such as nut and bolt arrangements, are inserted into the vanes to provide extra strength to hold the vanes under load.

A leading edge sheath is attached to the old edge to protect this edge.

The embodiments of the invention described above are merely examples, and many modifications will occur to those skilled in the art. It is therefore to be understood that this invention is not limited to the embodiments described in Section 2, but rather by the scope of the claims.

[Brief explanation of the drawing]

Fig. 1 is a diagram of a counter-rotating propeller device using the blades of the present invention; Fig. 2 is a diagram showing the airfoil portion of the front blade including the wing spar and coupling member; Fig. 3 is a diagram showing the front, rear, leading edge, Figure 4 shows the components of the trailing wing including the trailing edge, tip, root and spars; Figure 4 is a schematic representation of the wing spar showing the forward and aft undercut cavities and counterweights; Figure 5 The figure shows a pattern of blades having areas of constant thickness with respect to the cross section of the first shell, and Figure 6 shows the pattern of blades with areas of constant thickness with respect to the cross section of the second shell. Figure 7 is a schematic diagram illustrating the method of clamping the shell, spars, foam filler and adhesive in a mold to make the blade; Figure 8 is a diagram showing the vibration of the equipment; FIG. 9 is a diagram showing how the blades are balanced in the radial direction. FIG. 10 is a diagram showing how the blades are balanced in the chordal direction. Explanation of main symbols 302: Root part 307: Rear surface 309: Front surface 314: Front edge 316: Rear edge 318: IA girder

Claims (1)

[Claims] 1. A blade having first and second surfaces that intersect at the tip and root, a swept leading edge and a trailing edge, and a wing spar is made of a ductile, low-strength, low-modulus blade. In a method of making a composite material having unidirectional high modulus fibers placed in a matrix, a pattern representing a region of constant thickness with respect to the cross section of the blade is prepared for each side, and the composite material forming first and second shells in the pattern, aligning a wing spar between the first and second shells, and forming a cavity between the shells; , by placing the foam filler at the intended location and heating the vane for a suitable curing cycle for the low modulus matrix, the spar, the foam filler, the first shell and the second shell are heated. A method including the step of binding. 2. The step of bonding includes the step of applying an adhesive to the spar, the foam filler, and the inner surfaces of the first and second shells to maintain the structural integrity of the blade. Method described. 3. The method of claim 1 including the step of tightening the spar between the first and second shells so that the spar does not twist or separate the shells. 4. The method of claim 1, including the step of inserting a plurality of coupling members through the spar into the first shell and the second shell to restrain the spar to the shell. 5. inserting a plurality of coupling members into the first shell adjacent the leading and trailing edges of the spar to provide structural integrity between the spar and the shell; 5. The method of claim 4, comprising: 6. The method of claim 5, including the step of applying a leading edge protector along the leading edge of the vane. 7. In order to move the position of the geometric center of the blade, the blade includes a first counterweight and a second counterweight that move in the radial direction and the chord direction, and further includes a first and second counterweight. By adjusting a second counterweight, the center of the figure is positioned at a first predetermined position in the chord direction, and by modifying the first and second counterweights, the center of the figure is positioned 6. The method of claim 5, including the step of radially locating the second predetermined position. 8. Measure the chordwise position of the figure center for multiple blades, select the standard figure center position in the chord direction to minimize the scattering of the figure center position, and set the figure center of each blade to 8. The method of claim 7, including the step of locating at a standard graphic center position. 9. Measure the position of the figure center in the radial direction, and while keeping the position of the figure center at the standard figure center position in the chord direction, measure the position of the figure center in the radial direction to minimize scattering of the position of the figure center in the radial direction. Claim 8, further comprising the step of selecting a center position and locating the center of the figure for each of the plurality of blades at a standard center position of the figure in the radial direction.
Method described. 10. High setback, wide chord width made from a composite material with side-by-side parallel fibers with high modulus in one direction in a ductile, low strength, low modulus matrix. In a method of preserving the structural integrity of thin blades, first and second shells of the blade are provided, each shell representing one half of the blade, and shaped to the contour of the blade. having a plurality of composite layers of material in which the fibers within each layer are aligned in varying directions; applying an adhesive to the inner surface of the shell; bonding the spar, the first shell and the second shell by aligning the spar and curing the low modulus matrix in an autoclave or the like by a predetermined curing cycle for the matrix; method including. 11. The method of claim 10, including the step of tightening the shell to the spar to maintain the structural integrity of the blade. 12. The method of claim 11, wherein the tightening step includes inserting a plurality of coupling members into the first shell, the spar, and the second shell to couple the shell to the spar. 13. Claim 1 comprising passing a plurality of coupling members through the first shell and the second shell at a location adjacent to the leading edge of the wing spar.
The method described in 2. 14. Passing the plurality of coupling members through the first shell and the second shell at a location adjacent to the trailing edge of the wing spar.
Method described. 15. The method of claim 12, including the step of radially and chordally balancing the vanes. 16. The blade, which includes the first and second shell bodies composed of composite layers, the swept leading edge and the swept trailing edge, and the wing spar, which intersect at the wing tip and root portion, is made of ductile, low strength and low elasticity. In the method of making a plurality of composite layers, each having parallel unidirectional high modulus fibers placed in a matrix of modulus, a pattern is prepared for each layer, and the pattern is for each layer of the shell representing an area of constant thickness relative to the cross-section of the blade, and forming said composite layer in said pattern to form first and second shells each representing one half of the blade; molding and aligning a wing spar between the first and second shells and positioning a foam filler at a predetermined location to form a void between the shells; the wing spar, the foam filler; and applying adhesive to the inner surfaces of the first and second shells to maintain the structural integrity of the vane and heating the vane during a suitable curing cycle for the low modulus matrix. A method comprising bonding a spar, a first shell, and a second shell. 17. The method of claim 16, including the step of clamping the shell to the spar to maintain the structural integrity of the blade. 18. The method of claim 17, wherein the tightening step includes inserting a plurality of coupling members into the first shell, the spar, and the second shell to couple the shell to the spar. 19. The method of claim 18, including the step of inserting a plurality of coupling members into the first shell and the second shell adjacent the leading edge of the wing spar. 20. The method of claim 19, including the step of inserting a plurality of coupling members into the first shell and the second shell adjacent the trailing edge of the wing spar. 21. The method of claim 17, including the step of radially and chordally balancing the vanes. 22. Made from a composite material with parallel, unidirectional, high modulus fibers in a ductile, low strength, low modulus matrix, with intersecting fibers at the tip and root. a first and second surface having a swept leading edge and a swept trailing edge and a spar, the spar being movable radially and chordally for positioning the geometric center of the blade; In a method of balancing a blade having a counterweight and a second counterweight, by adjusting the first and second counterweights, the center of the figure is moved to the positioning the graphic center at a predetermined position in the radial direction by modifying the first and second counterweights. 23. Measure the position of the figure center in the chord direction for multiple blades, select the standard figure center position in the chord direction to minimize the scattering of the figure center position, and set the figure center for each blade. 23. The method of claim 22, including the step of locating at a standard graphic center position. 24. Measure the position of the figure center in the radial direction, and while keeping the figure center position at the standard figure center position in the chord direction, set the standard figure center in the radial direction to minimize the scattering of the figure center position in the radial direction. Claim 2, further comprising the step of selecting a position and locating the center of the figure for each of the plurality of blades at a standard center position of the figure in the radial direction.
The method described in 3.