Helicopter composite blade forming apparatus and method

Abstract

The present application belongs to the technical field of blade mould pressing forming technology in the helicopter composite blade manufacturing technology, and particularly relates to a helicopter composite blade forming device and method. The device comprises an upper tooling mould, a lower tooling mould and a blade skin support device arranged in a hollow spar. A blade hollow spar root positioning pin is fixed to the inner surface of one end of the lower tooling mould, and is used for positioning the blade root of the blade hollow spar during the forming process. A rear counterweight box body positioning pin is arranged near the other end of the lower tooling mould, and the positioning of the rear counterweight box body realizes the surface positioning of the end near the blade tip of the blade hollow spar. The upper tooling mould and the lower tooling mould are locked by a mould closing bolt. The process method adopted by the present application can solve the problems of uneven combination of the hollow spar and the blade skin, local area rich in resin, skin wrinkle and the like, and makes the formed product fully play the performance advantages.

Description

Technical Field

[0001] This invention belongs to the field of rotor blade compression molding technology in helicopter composite rotor blade manufacturing technology, specifically relating to a helicopter composite rotor blade molding device and method. Background Technology

[0002] The rotor system is a key component of a helicopter. The main rotor blades in the rotor system provide sufficient and stable lift for the helicopter as they rotate around the rotor shaft. The rotor system typically has 2 to 7 main rotor blades, and each main rotor blade needs to have the same and precise geometric shape and center of gravity position to ensure the stability of the conical surface formed when the blade rotates and minimize system vibration.

[0003] To ensure aerodynamic accuracy, helicopter composite rotor blades are typically manufactured using precise molding fixtures. In terms of process, composite rotor blades generally employ a co-curing method, where glass prepreg spar, foam core, and carbon fiber fabric prepreg are laid and positioned within the cavity of the lower mold of the molding fixture. Then, the upper mold of the fixture presses the entire rotor blade together, followed by heating and pressure curing using hot press equipment. Co-cured rotor blades have fewer components and fewer processing steps, allowing for a single molding process. During molding, the blade spars do not require pre-curing. However, during layup, the spars need to be manually shaped and positioned in the lower mold, which involves significant human intervention and requires a certain level of skill. As a result, the product quality stability is relatively low. For example, it is difficult to ensure the weight consistency of the blades when they are demolded, the center of gravity is prone to shift, and the static moment value needs to be adjusted in large amounts each time. During the curing process, the uncured spars inside the blades will be deformed by the interaction of the mold closing extrusion force, resulting in quality problems such as internal composite fiber deformation and wrinkling, and localized resin enrichment. Some of these problems can be detected by non-destructive methods such as X-ray and ultrasonic scanning, but some potential risks are not easy to detect by non-destructive methods.

[0004] Another method for manufacturing blades is through co-bonding molding. For example, the main load-bearing composite material components, such as the hollow spar of the blade, are pre-cured and pre-tested to ensure high quality consistency before blade molding. Alternatively, suitable components can be assembled through a selection process. During blade molding, the cured components are fixed in the mold according to specified positions, and adhesives such as adhesive films are applied between them. The adhesives cure and bond all the components together, completing the blade bonding process. This process easily ensures consistent blade quality, and the spar does not require a complex shape, resulting in a more uniform center of gravity. However, it also requires controlling many variables. For example, composite material spars are generally hollow, making manufacturing difficult and requiring precise shapes. The different thermal deformation coefficients between cured components, some uncured components, and molding tooling can easily generate interfacial residual stress. Detailed calculations of the thermal deformation of the component and tooling materials are needed to ensure matching and minimize internal stress. A typical hollow spar blade is shown below. Figure 1 .

[0005] Skin wrinkles and weak bonding issues: Compared to metal structures, the fiber orientation and thickness of hollow spars in composite material rotor blade structures can be optimized to a great extent to meet specific directional mechanical performance requirements. Composite material hollow spars are lightweight and high-strength, offering significant structural advantages and representing a more advanced design structure in modern helicopter rotor blades. However, manufacturing is relatively difficult, and the shape is prone to deviations of 0.2-0.5mm due to factors such as molding die materials, resin curing shrinkage, and pressure deformation. Furthermore, the degree of deviation varies depending on the geometric dimensions. For example... Figure 2 Geometric deviations can be localized or global. When shape deviations occur, uneven gaps and pressures will appear between the inner surface of the blade skin and the outer surface of the spar. As the curing temperature continues to rise, the adhesive in the bonding surface will flow to areas with larger gaps and lower pressure due to pressure. At the same time, the skin may develop wrinkles and weak bonding problems to varying degrees, requiring adjustments to the process to solve this problem.

[0006] The problem of high internal stress: Hollow spar blades typically have a counterweight chamber pre-installed inside the blade tip spars and rear foam cavity to adjust the static moment. This chamber also serves to position the spars, foam, and other components during blade curing. Once the spars are pre-cured, their geometry and length remain largely unchanged. The coefficient of thermal expansion of cured carbon fiber composite materials is typically 2.2e-6 / K, while the coefficient of thermal expansion of the steel used in blade forming molds is generally 1.2e-5 / K. This difference leads to a problem: the spars cannot elongate by the same amount as the mold due to thermal expansion. In this case, the positioning chamber for the blade spars will fail, resulting in significant stress between the skin and the spars. This issue needs to be addressed by improving the manufacturing process and adjusting the geometry. Summary of the Invention

[0007] The purpose of this invention is to provide a helicopter composite material rotor blade forming device and method, which focuses on solving problems such as skin wrinkling, weak bonding, large internal stress and positioning during the rotor blade curing process, and forms a relatively complete method for forming helicopter composite material hollow spar rotor blades.

[0008] The technical solution of the present invention: To achieve the above-mentioned objective, according to a first aspect of the present invention, a helicopter composite material rotor blade forming device is provided. The helicopter composite material rotor blade includes a hollow spar, a rear cavity filling material, a leading edge filling material, a rear counterweight box, a front counterweight box, an upper wing skin, and a lower wing skin. A hollow spar is disposed between the upper wing skin and the lower wing skin. The leading edge filling material is disposed at the outer front end of the hollow spar, the rear cavity filling material is disposed at the outer rear end of the hollow spar, and the rear counterweight box is disposed within the rear cavity filling material, with its side surface conforming to the rotor blade. The outer surface of the hollow spar; the front counterweight box is set inside the hollow spar of the blade; the forming device includes an upper tooling mold, a lower tooling mold, and a blade skin support device set inside the hollow spar; a locating pin for the root of the hollow spar of the blade is fixed on the inner surface of one end of the lower tooling mold, which is used to position the root of the hollow spar of the blade during the forming process; a locating pin for the rear counterweight box is opened at the other end near the lower tooling mold, and the locating of the hollow spar of the blade near the blade tip is achieved by the positioning of the rear counterweight box; the upper tooling mold and the lower tooling mold are locked together by mold closing bolts.

[0009] In one possible embodiment, the blade skin support device employs a flat silicone airbag.

[0010] In one possible embodiment, the upper and lower tooling molds are made of INVAR steel.

[0011] According to a second aspect of the present invention, a method for forming helicopter composite material rotor blades is provided, comprising the following steps: Step 1: Place the blade skin support device inside the hollow spar of the blade; Step 2: Perform surface pretreatment to improve the roughness of the outer surface of the hollow spar of the blade; Step 3: In sequence, lay down the lower airfoil skin, hollow blade spars, leading edge filling material, rear counterweight box, and rear cavity filling material on the inner surface of the tooling mold; locate the blade root of the hollow blade spars using the locating pin at the root of the hollow blade spars, and locate the rear counterweight box using the locating pin at the rear counterweight box. Step 5: Perform vacuum pretreatment on the lower mold of the tooling; Step 6: Place the upper wing skin inside the upper mold of the tooling, flip the upper mold of the tooling over, position it above the lower mold of the tooling, and lock it in place with the mold closing bolts; Step 7: Inflate the blade skin support device with gas to maintain its supported state; Step 8: Transfer the tooling as a whole to the curing equipment, and perform temperature-controlled curing of the entire blade according to the blade forming temperature control requirements; Step 9: Depressurize the blade skin support device, remove the blade skin support device, and remove the formed blade; Step 10: Open a front counterweight box mounting hole in front of the rear counterweight box on the molded blade, position the front counterweight box in the front counterweight box mounting hole using an adhesive clamp, and bond it by local heating and curing.

[0012] In one possible embodiment, in step 2, the surface pretreatment process includes removing the peelable fabric layer from the hollow spar of the blade or sanding the surface with sandpaper.

[0013] In one possible embodiment, in step 7, the air pressure of the blade skin support device is maintained at 0.4MPa-0.5MPa.

[0014] In one possible embodiment, in step 6, the mold closing gap is ≤0.1mm.

[0015] In one possible embodiment, in step 8, thermocouples are arranged in the high-temperature zone and low-temperature zone of the upper mold and lower mold of the tooling respectively to effectively control the temperature of the molding die.

[0016] Compared with the prior art, the beneficial effects of the present invention are as follows: 1) This invention is a hollow spar blade forming process method, which focuses on analyzing the technical points of forming blades with similar structures. Compared with co-curing blade forming technology, it is a brand-new blade forming solution. 2) The process method adopted in this invention can solve problems such as uneven bonding between hollow spar and blade skin, local areas of resin richness, and skin wrinkles, so that the molded product can fully exert its performance advantages. 3) The wing spar positioning method adopted in this invention can effectively solve the problem of large internal stress between the wing spar, skin and molding tooling, and avoid stress deformation and easy delamination of the blade after demolding. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of a hollow blade and hollow spar blade; wherein: 1 main blade, 1-1 blade skin, 1-3 hollow spar blade, 1-6 blade root, 1-7 blade tip, 1-8 front counterweight cavity box, 1-9 rear counterweight cavity box, 1-3-1 hollow spar blade / blade root bushing. Figure 2This is a cross-sectional schematic diagram of the blades and tooling; where: 1-2 is the lower wing skin of the blades, 1-4 is the leading edge filler of the blades, 1-5 is the foam filler of the blades, 2 is the blade forming tooling, 2-1 is the upper mold of the blade forming tooling, and 2-2 is the lower mold of the blade forming tooling. Figure 3 This is a schematic diagram of the positioning of the hollow blade spar and the tooling root; where: 2-2-1 is the positioning pin at the root of the hollow blade spar, 1-3-1 is the mounting hole for the bushing of the blade and hub assembly, 1-3-2 is the cutting line for the blade process allowance, 1-3-3 is the positioning pin hole (process hole) for the hollow blade spar, and 1-3-4 is the blade process allowance area. Figure 4 A schematic diagram of the hollow spar and blade skin support device for the propeller blade; where: 1-3-5 is the sealing device at the tip of the hollow spar, and 2-3 is the silicone airbag; Figure 5 A schematic diagram showing the positioning of the hollow spar and counterweight box of the propeller blade; where: 1-10 are reinforcing ribs in the foam area of ​​the propeller blade; Figure 6 The diagram shows the blade curing parameters; where: T1 is room temperature, T2 is the resin flow isothermal temperature, T3 is the resin curing isothermal temperature, T4 is the demolding temperature, t1 is the time to enter the resin flow isothermal temperature, t2 is the time to exit the resin flow isothermal temperature, t3 is the time to enter the curing isothermal temperature, t4 is the time to exit the curing isothermal temperature, t5 is the demolding time, and P is the curve of the pressure intervention of the hollow spar and blade skin support of the blade. Figure 7 This is a schematic diagram of the wingtip front counterweight assembly structure; where: 1-3-6 is the front counterweight box assembly hole on the surface of the hollow spar of the rotor blade, and 1-3-7 is the adhesive bonding surface between the front counterweight box and the hollow spar of the rotor blade and the skin. Detailed Implementation

[0018] To further understand the present invention, preferred embodiments of the present invention are described below in conjunction with examples. However, it should be understood that these descriptions are only for further illustrating the features and advantages of the present invention and not for limiting the claims of the present invention.

[0019] Hollow blade spars are generally made of high-temperature composite materials. To ensure structural stability and withstand the secondary heating during blade forming, the curing temperature of hollow blade spars is generally above 180℃.

[0020] Example 1 A helicopter composite material rotor blade forming device is disclosed. The helicopter composite material rotor blade includes a hollow spar, a rear cavity filling material, a leading edge filling material, a rear counterweight box, a front counterweight box, an upper wing skin, and a lower wing skin. A hollow spar is disposed between the upper and lower wing skins. The leading edge filling material is disposed at the outer front end of the hollow spar, the rear cavity filling material is disposed at the outer rear end of the hollow spar, and the rear counterweight box is disposed within the rear cavity filling material, its side surface conforming to the outer surface of the hollow spar. The front counterweight box is disposed within... The forming device includes an upper tooling mold, a lower tooling mold, and a blade skin support device disposed within the hollow blade spar. A locating pin for the root of the hollow blade spar is fixed to the inner surface of one end of the lower tooling mold for positioning the root of the hollow blade spar during the forming process. A locating pin for a rear counterweight box is provided at the other end near the lower tooling mold. The locating pin for the rear counterweight box is used to position the hollow blade spar near the blade tip. The upper tooling mold and the lower tooling mold are locked together by mold-closing bolts.

[0021] like Figure 3 As shown, the hollow spar of the rotor blade requires a pre-treatment process before blade forming. Precise positioning holes are made at the root of the hollow spar to facilitate static moment, torsion, and shape inspections. This also allows for precise positioning of the spar root during blade forming. These positioning holes can be process holes independent of the hollow spar's manufacturing allowance area, or they can utilize the bushing holes connecting the hollow spar to the rotor hub. However, they cannot be final state holes to avoid damage.

[0022] Before use, the bonding area between the hollow blade spars and the blade skin needs to be roughened. This can be done by grinding, but to avoid damaging the fiber layer, a peelable fabric layer is laid on the surface of the hollow blade spars during molding. The texture of the fabric layer effectively increases the surface roughness of the bonding surface. The peelable fabric is removed before bonding.

[0023] refer to Figure 4To ensure a proper fit between the hollow blade spars and the blade skin, an outward supporting force needs to be applied to the inner surface of the hollow blade spars during use. One function of this supporting force is to induce slight deformation in the cured hollow blade spars, ensuring full contact with the blade skin. Another function is to provide sufficient reverse supporting force to the hollow blade spars when the blade forming fixture applies pressure, preventing secondary deformation of the hollow blade spars during the extrusion of the skin. The supporting force can be provided by the mechanical fixtures, but the torsional changes in the shape within the beam are detrimental to the pressure applied by the mechanical fixtures, weakening the supporting effect. The supporting force can also be provided by pressurized air bags such as vacuum bags or nylon ducts. This invention employs a special silicone airbag structure, which can be easily inserted into the hollow beam and is not easily punctured by the cured composite material beam during pressurization. The inflation pressure is adjusted according to the actual wall thickness of the hollow blade spars and the strength of the fixture; the pressure provided in this invention is 0.4MPa-0.5MPa.

[0024] refer to Figure 5 The blade forming fixture mainly consists of two parts: an upper mold (not shown) and a lower mold, connected by bolts and nuts. During layup, the bolts and nuts are loosened to remove the upper mold. The placement and positioning of the main materials are performed on the lower mold.

[0025] The blade skin and leading-edge spars are laid on the lower mold surface. The hollow blade spars are positioned in the lower mold cavity through the root process holes. The double process bushing holes at the root position position the hollow blade spars in both the angular and spanwise (length) directions, respectively, and also determine the chordal position of the hollow blade spars at the root. The hollow blade spars in the wingtip area are not forcibly positioned; they are aligned and placed using auxiliary positioning fixtures (not shown), and the chordal position is initially determined. The advantage of this positioning method is that it minimizes the inconsistency in elongation between the spars, skin, and fixtures caused by temperature changes during the molding and heating process, preventing damage to the hollow blade spars due to molding and reducing internal stress in the skin. This process can also be improved by adjusting the fixture material, such as changing the main body of the fixture to INVAR steel, which undoubtedly increases manufacturing costs.

[0026] After the hollow spar of the blade is positioned, the rear cavity filling material and trailing edge reinforcement layer are laid. The filling material is generally low-density PMI foam or aramid honeycomb material, each with its own advantages and disadvantages depending on the specific needs. Foam materials are easier to process into complex geometries, but have lower shear and peel strength and poorer impact resistance. Honeycomb materials have better impact resistance, shear strength, and peel strength than foam materials, but have a more pronounced edge effect and are not easily processed into complex geometries. The choice depends on the specific conditions of the required location. The trailing edge reinforcement layer is laid in the trailing edge region of the blade to improve its torsional stiffness.

[0027] The blade tip employs a pre-installed counterweight box, consisting of a front counterweight box and a rear counterweight box, used to install counterweight plates for adjusting the static moment and dynamic balance parameters of the blade. Typically, the front counterweight box is located within the hollow spar of the hollow blade, while the rear counterweight box is located within a mandrel outside the hollow spar. However, once the counterweight cavities are pre-installed within the blade, the front and rear counterweight boxes will experience inconsistent spanwise displacements due to the varying thermal deformation coefficients between the hollow spar and the tooling, making it impossible to guarantee chordal alignment. In this case, refer to... Figure 5 The positioning requirements for the hollow spar of the rotor blade necessitate adjusting the counterweight box within the hollow spar for subsequent assembly. Simultaneously, the installation of the rear counterweight facilitates the positioning of the hollow spar during molding, ensuring it does not rotate in the chord direction.

[0028] refer to Figure 6 After the blades are molded, a curing process is required. The curing parameters should be referenced from the curing parameters of the selected raw materials. If the curing parameters given in various material standards differ slightly, or when the product size is large, the setting range of curing parameters and the heating rate need to be comprehensively considered to ensure temperature uniformity. Other influencing factors may include analysis and control in the fields of Dynamic Mechanical Analysis (DMA), Thermogravimetric Analysis (TGA), and Differential Scanning Calorimetry (DSC). This ensures that the performance of the cured product meets the usage requirements. All connection interfaces, especially the connection interface between the blade skin and the hollow blade spars, should be tested with furnace-fired specimens of the same mechanical performance requirements as the molded state to objectively verify the bonding quality stability inside the blades.

[0029] After the blades have solidified and formed, they need to be cleaned and cut to remove excess material. The weight after demolding is recorded and compared with the weight upon entering the mold, serving as one of the objective bases for weight adjustment.

[0030] refer to Figure 7 At the blade tip, on the same cross-section of the hollow spar surface as the rear counterweight box, a mounting hole for the front counterweight box is created. The pre-fabricated front counterweight box is then glued into this mounting hole. While using glue, screws can be added to further enhance the assembly strength of the counterweight box. This process ensures that the front and rear counterweights are positioned uniformly, unaffected by thermal deformation, and facilitates adjustments to the static torque and dynamic balance.

[0031] Example 2: 1) Clean the hollow spar 1-3 of the blade after curing, remove the process allowance, and make process holes 1-3-3.

[0032] 2) Reference Figure 4High-temperature adhesive film is used to attach the sealing device 1-3-5 at the tip of the hollow spar of the blade to the tip of the hollow spar of the blade. The device is positioned by pressure using a vacuum bag method or tooling positioning method, and then bonded and cured at a high temperature of 170℃.

[0033] 3) Remove the peelable cloth / sandpaper from the hollow spar 1-3 of the blade, place the blade skin support device 2-3 inside the hollow spar of the blade, preheat the surface of the hollow spar of the blade with a hot air gun, and apply adhesive film for later use.

[0034] 4) Reference Figure 5 Open the blade forming fixture 2. The lower mold 2-2 is used for positioning the main material layup, and the upper mold 2-1 is used to lay the upper wing skin.

[0035] 5) In sequence, lay down the wing skin, leading edge filler 1-4, blade hollow spar 1-3, rear counterweight box 1-9, rear cavity filler foam 1-5 and other parts on the surface of the lower mold 2-2, and perform vacuum pretreatment.

[0036] 6) The hollow spar 1-3 of the blade is positioned at the blade root 1-6 within the tooling positioning pin 2-2-1 via process hole 1-3-3, and the tip of the hollow spar is positioned via the side elevation of the rear counterweight box 1-9. The double process bushing holes at the root position position the hollow spar in both the angular and spanwise (lengthwise) directions, and also determine the chordal position of the root of the hollow spar.

[0037] 7) Remove the upper mold skin from the tooling and place it on the upper surface of the pre-positioned propeller parts.

[0038] 8) Flip the upper mold of the tooling and position it above the lower mold of the tooling. Tighten it with the mold closing bolts. The mold closing gap needs to be ≤0.1mm.

[0039] 9) Transfer the entire tooling to the curing equipment and connect the temperature control thermocouples. Typically, thermocouples should be placed in both the high-temperature and low-temperature zones of the tooling to effectively control the temperature of the molding die.

[0040] 10) Check the airtightness of the hollow spar and blade skin support device (silicone airbag) 2-3.

[0041] 11) Perform the curing process according to the characteristics of the raw materials. For example: Set the constant temperature T1 for the first stage to 80℃ for 2 hours; set the constant temperature T3 for the second stage to 130℃ for 3 hours. Before heating, set the pressure P of the silicone airbags 2-3 to 0.5MPa. After completing the second constant temperature plateau T3, begin cooling. To reduce the stress level between the tooling and the blades after curing, depressurize the tooling typically between 60℃ and 80℃.

[0042] 12) Open the mold and remove the blade.

[0043] 13) After the blade is formed, the blade is precisely positioned and tested through the root process hole 1-3-3 and the rear counterweight cavity box 1-9. At the same time, the blade and hub assembly bushing mounting hole 1-3-1 and the front counterweight box 1-8 mounting hole 1-3-6 are also installed.

[0044] 14) Position the front counterweight box 1-8 in the mounting hole 1-3-6 using a bonding clamp, and perform local heating and curing bonding.

[0045] 15) After the blades pass inspection, they are cut according to the blade process allowance cutting line 1-3-2 to form the final molded blade product.

Claims

1. A method for forming composite material rotor blades for helicopters, characterized in that, For forming hollow spar blades, the forming device includes an upper tooling mold, a lower tooling mold, and a blade skin support device disposed within the hollow spar. A locating pin for the root of the hollow spar blade is fixed to the inner surface of one end of the lower tooling mold for positioning the root of the hollow spar blade during the forming process. A locating pin for a rear counterweight box is provided at the other end near the lower tooling mold, enabling positioning of the hollow spar blade near the blade tip. The upper and lower tooling molds are locked together by clamping bolts. The molding method includes the following steps: Step 1: Place the blade skin support device inside the hollow spar of the blade; Step 2: Perform surface pretreatment to improve the roughness of the outer surface of the hollow spar of the blade; Step 3: In sequence, lay down the lower airfoil skin, hollow blade spars, leading edge filling material, rear counterweight box, and rear cavity filling material on the inner surface of the tooling lower mold. Position the blade root of the hollow blade spars using the locating pin at the root of the hollow blade spars, and position the rear counterweight box using the locating pin at the rear counterweight box. Position the tip of the hollow blade spars using the side elevation of the rear counterweight box. Step 4: Perform vacuum pretreatment on the lower mold of the tooling; Step 5: Place the upper wing skin inside the upper mold of the tooling, flip the upper mold of the tooling over, position it above the lower mold of the tooling, and lock it in place with the mold closing bolts; Step 6: Inflate the blade skin support device with gas to maintain its supported state; Step 7: Transfer the tooling as a whole to the curing equipment, and perform temperature-controlled curing of the entire blade according to the blade forming temperature control requirements; Step 8: Depressurize the blade skin support device, remove the blade skin support device, and remove the formed blade; Step 9: Open a front counterweight box mounting hole in front of the rear counterweight box on the molded blade, position the front counterweight box in the front counterweight box mounting hole using an adhesive clamp, and bond it by local heating and curing.

2. The method for forming helicopter composite material rotor blades according to claim 1, characterized in that, The blade skin support device uses a flat silicone airbag.

3. The method for forming helicopter composite material rotor blades according to claim 1, characterized in that, The upper and lower molds of the tooling are made of INVAR steel.

4. The method for forming helicopter composite material rotor blades according to claim 1, characterized in that, In step 2, the surface pretreatment process includes removing the peelable fabric layer from the hollow spar of the blade or sanding the surface with sandpaper.

5. The method for forming helicopter composite material rotor blades according to claim 1, characterized in that, In step 7, the air pressure of the blade skin support device is maintained at 0.4MPa-0.5MPa.

6. The method for forming helicopter composite material rotor blades according to claim 1, characterized in that, In step 6, the mold closing gap is ≤0.1mm.

7. The method for forming a helicopter composite material rotor blade according to claim 1, characterized in that, In step 8, thermocouples are arranged in the high-temperature zone and low-temperature zone of the upper mold and lower mold respectively to effectively control the temperature of the molding mold.

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