Engine blade with integrated heating element

Abstract

The present invention relates to an engine blade (20) for an aircraft engine (100), the engine blade (20) comprising a blade body (22) having a blade surface (23) and heating elements (40, 40A, 40B) for heating at least a portion of the blade surface (23). The heating elements (40, 40A, 40B) are integrally provided with the blade body (22) within the blade surface (23). The present invention further relates to a method for manufacturing the engine blade (20), wherein composite layers (321, 322, 323, 324) are provided for the engine blade (20), the heating elements (40, 40A, 40B) are provided and applied on at least one of the composite layers (340), and the composite layers (321, 322, 323, 324, 340) for the engine blade (20) are laminated and bonded to form the engine blade (20).

Description

Detailed Description of the Invention

[0001] The present invention relates to engine blades and engine rotors for aircraft. The present invention also relates to a method of forming the engine blades and engine rotors.

[0002] It is known that rotors are used for aircraft fan engines. Those rotors are typically made from metallic materials in order to provide durability and mechanical stability. Also, in order to reduce the weight of such rotors and thereby reduce the overall weight of the aircraft equipped with each, it is known to explore lighter materials. One option for such lightweight materials is a so-called composite material comprising a polymer matrix material and reinforcing fibers. In aircraft, usually only one type of such polymer matrix and fiber combination is used. In particular, the matrix, which is the polymer matrix, embeds the reinforcing fibers, and the reinforcing fibers are continuous fibers or so-called endless fibers. The use of such continuous fibers allows for a specific orientation of such fibers within the polymer matrix during the manufacturing process. This leads to a well-known construction in which the direction and orientation of those continuous fibers within each polymerized body portion of the rotor blade are well known. Since the mechanical stability and the possibility of providing a load path for the loads applied to such rotor blades are strongly correlated with the orientation of the reinforcing fibers, knowledge about such orientation is important for calculating the stability and thus for constructing the rotor blades accordingly. Thus, in the generally known efforts to provide lightweight rotor blades, such rotor blades can only be produced by such continuous fibers as reinforcing fibers within the matrix material.

[0003] The problem with the current situation is that applying endless or continuous fibers into the matrix material is relatively easy for simple rotor blade shapes. However, with respect to more complex structures, endless fibers present challenges, particularly in the root section where the blades are attached to the shaft, or the complexity of the production method is greatly increased. Therefore, according to current knowledge, these composite rotor blades are used in combination with a conventional shaft hub (for example, made from steel) to provide a connection to the drive shaft and support the rotor blades when bearing the drive load.

[0004] A further disadvantage of the current solution is the fact that the rotor blades not only need to provide stability to withstand the driving load, but also need to withstand additional reaction loads resulting from air and centrifugal loads during rotation, as well as from mechanical impacts coming from birds or stones, etc. To provide this additional or secondary mechanical stability, the rotor blades are constructed with increased size and, among other things, increased weight.

[0005] A further disadvantage of the current situation is that ice can easily form on the rotor blades. For example, when the room temperature is low but the relative humidity of the ambient air is high, there is a high risk of ice forming on the rotor blades. As a result of ice formation, the rotor blades and rotor change their dynamic and aerodynamic properties, respectively. This can lead to several disadvantages, such as an undesirable weight increase in the aircraft or rotor imbalance that leads to undesirable vibrations of the aircraft. Both weight increase and imbalance can cause a loss of efficiency and, therefore, a loss of range. Moreover, the current situation does not allow for the removal of ice formed during flight. Furthermore, currently, the formed ice cannot be removed in a timely and / or efficient manner.

[0006] The object of the present invention is to solve the above-mentioned problems at least partially.

[0007] In particular, an object of the present invention is to provide a simple and cost-effective method for reducing the weight of a composite material rotor while still maintaining high mechanical stability. A further object of the present invention is to provide a lightweight rotor with a heating function for de-icing its blades.

[0008] At least some of the described objectives are achieved by the engine blade described in claim 1, the engine rotor described in claim 8, the propulsion unit described in claim 9, the method for manufacturing the engine blade described in claims 11 and 12, and the method for manufacturing the engine rotor assembly described in claim 15.

[0009] Further advantages and features of the present invention can be derived from the dependent claims, description, and drawings. Features and details described in relation to the engine blade of the present invention naturally also apply in relation to the engine rotor, propulsion unit, and method of the present invention, and vice versa, and references are always made to and / or can be made to each other with respect to the disclosure of individual embodiments of the present invention.

[0010] The present invention provides an engine blade for an aircraft engine. The engine blade includes a blade body, the blade body including a blade surface and a heating element for heating at least a portion of the blade surface. The heating element is provided integrally with the blade body within the blade surface.

[0011] In other words, an engine blade for an aircraft engine is disclosed, which has a blade body, the blade body integrally provided with a heating element for heating at least a portion of the surface of the engine blade.

[0012] In this context, engine blades can preferably be understood as wings that exert thrust on the air when rotated around a rotation axis. Aircraft engines can be electric engines, which drive the engine blades to rotate.

[0013] The engine blade includes a blade body having a blade surface, where the blade surface can preferably be understood as an outer surface surrounding the blade body.

[0014] The blade body further includes a heating element, which can preferably be understood as a device or structure for supplying de-icing energy. For example, the heating element can be a device or structure for releasing heat. For example, the heating element can transport a substance containing thermal energy, or it can generate heat by converting supplied energy into thermal energy. Alternatively, instead of providing thermal energy, the heating element could convert the supplied energy into other forms of energy, such as vibrational (ultrasonic, high-frequency) energy for removing ice from the blade surface.

[0015] The heating element is provided integrally with the blade body within the blade surface.

[0016] In this context, an integrated configuration can preferably be understood as consisting of only one piece. For example, the heating element can be contained within the blade body. Preferably, in an integrated configuration, different components or materials of the blade body can be connected to one another through interlocking connections or joints, for example, the parts are joined together by fusion, melting, and / or intermolecular or chemical bonding forces, sometimes using additives (e.g., adhesives).

[0017] Furthermore, the expression "within the blade surface" can preferably be understood as being below or above the blade surface. Thus, the heating element can be located inside the blade body and / or at least partially surrounded by the blade surface. Alternatively, the heating element can also form at least a portion of the blade surface.

[0018] This makes it possible to provide heating functionality to the engine blades while maintaining structural integrity without adding a significant amount of additional weight, and also simplifies the manufacturing and maintenance of the corresponding aircraft engines.

[0019] Advantageously, the heating element may preferably include at least one heat transfer conduit. Preferably, the heat transfer conduit may be electrically conductive. Alternatively or additionally, the heat transfer conduit may extend along a winding (curved) extension path. Alternatively or additionally, the heat transfer conduit may extend in conjunction with the blade surface. Preferably, the heat transfer conduit may function as an electric resistance heater. Alternatively or additionally, the heat transfer conduit may form a passage for transporting a heat transfer medium. Alternatively or additionally, the heat transfer conduit may be thermally conductive and capable of transferring thermal energy through heat conduction.

[0020] This makes it possible to provide pathways for releasing or supplying de-icing energy. In particular, electrically conductive pathways may be effective in generating enough heat to de-ic the blade surface.

[0021] Advantageously, the heating element can preferably be provided on a portion of the blade body in an additive manufacturing method. For example, the heating element can be provided by 3D printing. Alternatively or additionally, the heating element can be provided in a deposition method.

[0022] This allows the heating element to be provided directly on top of the blade body. Furthermore, the heating element can be provided as a microstructure. This allows the de-icing function to be integrated into the engine blade without compromising the structural integrity or weight of the engine blade. Moreover, the size and shape of the heating element can be freely selected, and the manufacturing speed can be increased.

[0023] Advantageously, the blade body can be made from a composite material, preferably comprising a polymer matrix and reinforcing continuous fibers embedded within the polymer matrix. Advantageously, the blade body can preferably include a layered structure, which can preferably be a composite structure. The layered structure can include a base layer for transmitting driving force, which can be configured to transmit driving force from an aircraft engine and exert thrust through the blade surface into the air. The layered structure can further include a heating layer containing a heating element.

[0024] Thus, the weight-related advantages of composite structures can be advantageously combined with the benefit of providing de-icing functionality within the engine blade. For example, the layers of the structure can include matrix material and reinforcing components. For example, the layered structure can include resin, polymer, and / or glass fiber. The layers can have a material thickness of 0.1 mm to 10 mm.

[0025] Advantageously, the heating layer can preferably be positioned between the base layer and a cover layer that covers the heating layer on the opposite side of the base layer.

[0026] As a result, the cover layer can be used not only to provide a protective cover for the heating element, but also to form a material bond between the base layer and the heating layer. Consequently, the heating element can be protected from damage during flight.

[0027] Advantageously, the heating layer may include a carrier layer. For example, the carrier layer may be a foil or a film. The heating layer may further include a heating element, which may be provided integrally with the carrier layer or integrally on top of the carrier layer. For example, the heating element may be printed or deposited on top of the carrier layer. Advantageously, the heating layer may be integrally bonded to the base layer via the carrier layer by a material bond.

[0028] This allows for increased reliability and design flexibility of the heating element, because the heating element can be manufactured and tested independently before being integrated into the blade body.

[0029] Advantageously, the heating layer can preferably form an integral part of the base layer. For example, the base layer can include at least two base support layers. The heating layer can be disposed between the base support layers, or the heating layer can be disposed on one of the base support layers. It is also conceivable that the aforementioned carrier layer can be one of the base support layers.

[0030] Thereby, the heating layer contributes to the mechanical strength of the engine blade. Thus, the engine blade can be provided with deicing functionality without reducing the robustness and strength of the engine blade.

[0031] Advantageously, the heating element can preferably be disposed on the radially outward portion of the engine blade. Preferably, the heating element can be disposed between the tip section and the root section of the blade body. Moreover, the engine blade can include two or more heating elements. For example, two heating elements can extend on opposite sides of the blade body. Alternatively or additionally, one or more heating elements can extend or be distributed to cover the entire blade surface for heating.

[0032] Thereby, the deicing functionality can be provided in various sections of the engine blade that are typically most affected by ice formation during flight.

[0033] Advantageously, the heating element can include a connection section for supplying a heat generation agent. Preferably, the heating element can extend between the tip section of the blade body and the root section for attaching the engine blade to the drive shaft. Preferably, the connection section can be provided in the root section. For example, the connection section can be an electrical connector.

[0034] The present invention provides an engine rotor for propelling an aircraft. The engine rotor includes at least one of the engine blades. For example, the aircraft may be an aircraft for vertical takeoff and landing. The engine rotor may include a hub device and a plurality of engine blades, the plurality of engine blades may be arranged to extend radially from the periphery of the hub device. Preferably, the engine blades may be set around the periphery of the hub device at a pitch to form a helical spiral.

[0035] In this context, the engine rotor (which may also be referred to as the rotor hereafter) can preferably be understood as a drive device for an aircraft engine, and it includes engine blades and rotates with the engine blades to generate thrust. The hub device can be the central part of the engine rotor, and all the engine blades can be connected to it. Preferably, the hub device can be a ring. Preferably, the hub device can be connected to the drive shaft of the aircraft engine.

[0036] The present invention provides a propulsion unit for propelling an aircraft. The propulsion unit includes an engine rotor and an engine housing equipped with an engine duct, wherein the engine rotor is rotatably disposed within the engine housing about an axis of rotation.

[0037] Further embodiments of the present invention described above can derive the same advantages and technical benefits with respect to engine blades as those described above.

[0038] Advantageously, the engine rotor and / or propulsion unit may include a factor supply device for supplying a heat-generating factor to a heating element for heating the blade surface. Preferably, the factor supply device is capable of supplying a heat-generating factor to a connection section of the heating element. Preferably, the factor supply device is capable of transmitting the heat-generating factor via cable coupling (e.g., through a wire connection) or non-contact (e.g., wirelessly).

[0039] In one example, the action factor supply device may preferably include an inductive supply device for inductively supplying electrical energy to a heating element as a heat generation action factor. Therein, the engine housing may include the inductive supply device. The inductive supply device may preferably be located in the upstream section of the engine duct. The inductive supply device may preferably be located upstream of the engine blades. Furthermore, the inductive supply device may preferably be located radially outward from the engine blades. Preferably, the heating element may correspondingly be provided as an electrical resonant circuit for inductively receiving electrical energy from the inductive supply device.

[0040] This allows the energy required for de-icing to be transmitted wirelessly to the engine blades. Consequently, cables and electrical connections can be avoided, reducing the structural complexity of the propulsion unit.

[0041] Alternatively or additionally, the action factor supply device may preferably include a slip ring for supplying electrical energy to a heating element as a heat generation action factor. Therein, the engine rotor may include the slip ring. Preferably, the slip ring may be integral with the engine rotor. Furthermore, the slip ring may be positioned to be movable relative to the engine housing. Preferably, the slip ring may be positioned coaxially with the axis of rotation, and preferably radially inward with respect to the root section.

[0042] In this context, the slip ring can preferably be understood as a power transmission device for transferring power from a stationary structure (e.g., engine housing) to a rotating structure (e.g., engine rotor).

[0043] This allows the engine rotor to be integrally equipped with means for supplying electrical energy to the heating element, even though the engine blades are rotating during flight.

[0044] According to the present invention, an aircraft for vertical takeoff and landing is provided. The aircraft includes at least one of the propulsion units.

[0045] The present invention provides a method for manufacturing an engine blade having a blade body with an integrated heating element for heating at least a portion of the blade surface. Therein, the heating element can be printed or deposited on the blade body. Furthermore, the heating element can be covered with a cover layer to form at least a portion of the blade surface to be heated. For example, the cover layer can be an adhesive or a resin.

[0046] For example, the heating element can be provided in a process suitable for generating a three-dimensional structure (e.g., etching, vapor deposition, or printing).

[0047] This allows the heating element to be provided directly on the blade body, without altering the structural integrity of the blade body. A cover layer can provide a protective cover for the heating element. This not only allows for provision to the engine blade in a minimum number of manufacturing steps, but it also facilitates the easy integration and testing of the de-icing functionality onto the engine blade. Furthermore, the heating element can be provided in very close proximity to the blade surface, thus improving de-icing efficiency. This is because the released energy only needs to travel through fewer layers.

[0048] The present invention provides a method for manufacturing an engine blade having a blade body portion equipped with an integrated heating element for heating at least a portion of the blade surface. A composite layer for the engine blade is provided. The heating element is provided and applied on at least one of the composite layers. The composite layers for the engine blade are laminated. The composite layers are joined together to form an engine blade.

[0049] Advantageously, providing and applying the heating element can preferably be done before lamination. Preferably, the composite layer can be an uncured layer. Advantageously, providing and applying the heating element can include, for example, printing the heating element on at least a portion of the composite layer. Alternatively or additionally, it can include directly printing at least one electrically conductive heat transfer conduit on the composite layer. Alternatively or additionally, it can include arranging and curing the heating element (preferably together with a carrier layer) on at least a portion of the composite layer. Alternatively or additionally, it can include arranging and bonding the heating element (preferably together with a carrier layer) on at least a portion of the composite layer.

[0050] In any of the above configurations, the heating element can be provided on top of a structural carrier element (e.g., a fiberglass-reinforced slat), which can then be integrated into the blade body to form its respective parts, thereby contributing to force transmission after all layers are joined together. Consequently, the mechanical strength of the engine blade can be improved. An additional benefit is that the manufacturing speed can be increased because the configuration of this method facilitates the provision of the heating element in a step independent of providing the other composite layers.

[0051] Advantageously, bonding the composite layers can preferably include molding, injection molding, composite pressing, lamination, curing, and / or bonding.

[0052] For example, pre-cut reinforcing layers can be repeatedly placed into a mold and covered with resin. Once the desired structure is realized, curing can be initiated.

[0053] The present invention provides a method for manufacturing an integrated engine rotor assembly. The method provides a plurality of engine blades, each including a blade body with a blade surface. Therein, at least one of the engine blades is manufactured in one of the above methods for manufacturing an engine blade. Furthermore, a hub device is provided for receiving and supporting the root section of the engine blade. The engine blade and the hub device are arranged to form an engine rotor assembly. The engine blade and the hub device are joined together to form an integrated engine rotor assembly.

[0054] In this context, the engine rotor assembly can preferably be understood as a structure comprising a combination of different parts (for example, at least the engine rotor).

[0055] With the above configuration, all parts typically involved in an engine rotor assembly can be supplied as a single integrated unit. This allows a single, unified structure to be received as the engine rotor assembly at the end of the process.

[0056] Advantageously, joining the engine blades and hub devices can preferably involve injection molding, compound pressing, and / or bonding.

[0057] Advantageously, providing a hub device may preferably involve forming the hub device by pressing or injection molding. For example, the hub device may be made from a composite material comprising a polymer matrix and shredded reinforcing fibers embedded within the polymer matrix. Preferably, the step of forming the hub device may include providing and positioning an action factor supply device in the hub device in order to provide an action factor supply device integrated with the hub device.

[0058] This allows other functional components (which may be needed to supply energy to the heating element for de-icing) to be easily embedded in the engine rotor assembly, and an integrated engine rotor, including the heating element and action factor supply device within the engine blades, can be received in a single manufacturing process. The advantage of using shredded fibers is that they result in an isotropic material, which allows them to bear mechanical strain or load equally well regardless of its direction.

[0059] Alternatively or additionally, the present invention discloses further configurations of engine rotors (hereinafter simply referred to as “rotors”). Thus, features described below relating to rotors, engines, or methods are equally applicable to the aforementioned embodiments of the present invention. For example, the features of the “rotor blades” described below are equally applicable to the aforementioned “engine blades.”

[0060] According to the present invention, a rotor is constructed for use in a ducted fan engine. Such a rotor includes a plurality of rotor blades, each having a blade body, the blade body comprising a root section, the root section being connected to a drive shaft and intended to receive primary drive loads and provide a primary load path. Additionally, such a blade body comprises a tip section, the tip section being positioned at the opposite end of the blade body. The plurality of rotor blades are arranged circumferentially to form the rotor described above. Additionally, the rotor includes a rotationally symmetric load bearing element, the rotationally symmetric load bearing element in force-transmitting contact with the plurality of blade bodies in the load bearing section. Such a load bearing section is positioned between all the root and tip sections of the blade body, and therefore the load bearing element is capable of providing a secondary load path for at least a portion of the reaction loads of the blade body. Furthermore, the blade body of the rotor blade is made from a composite material comprising a polymer matrix and reinforcing continuous fibers embedded within the polymer matrix. The load-bearing element is also made from a composite material comprising a polymer matrix, which further comprises shredded reinforcing fibers embedded within such polymer matrix.

[0061] According to the present invention, the rotor is made of a composite material, thereby offering the possibility of lightweight construction and, respectively, the possibility of reducing the overall weight when used in an aircraft. To provide the stability of the rotor blades as the primary functional element itself, the composite material for their blade bodies comprises a polymer matrix with known reinforcing continuous fibers. This means that the blade bodies can be constructed in a generally known manner, thereby providing the necessary known mechanical stability to withstand the driving load from the rotating shaft.

[0062] In contrast to generally known techniques, the mechanical stability against secondary loads (i.e., reaction loads applied to the rotor blades) is, here, at least partially, shifted from the blade body to secondary elements formed as load-bearing elements according to the present invention. Thus, such load-bearing elements are separate from the connection to the shaft, thereby providing a separate so-called secondary load path for at least partially receiving those secondary loads, such as reaction forces coming from the blade body. Due to the fact that such load-bearing elements are in force-transmission contact with all of the blade body, they are capable of receiving reaction loads from all of those blade body sections. Furthermore, the force-transmission contact is located between the tip section and the root section, so that a lever effect is generated between the reaction force at the tip of the rotor blade and its respective transmission into the secondary load path in the intermediate load-bearing section of the blade body. The same applies to any impact that results in the respective reaction forces at the central hub where the shaft is connected to the rotor blades.

[0063] According to the above discussion, the load-bearing section is therefore located in the intermediate portion between the tip section and the root section of the blade body, and is spaced apart from the tip section and the root section of the blade body. For example, the load-bearing element can form a kind of symmetrical ring that surrounds the drive shaft, thereby connecting all of the blade body in this circumferential extension, and thereby connecting the rotor blades. As will be discussed later, the load-bearing element can provide further functionality, in particular, to improve the aerodynamic functionality of the rotor itself.

[0064] According to the present invention, one objective is to combine high mechanical stability with maximum weight reduction. To achieve this objective, the present invention provides load-bearing elements made of composite materials, for example, because they are lightweight materials in themselves. Here, due to the fact that the load-bearing element separates the secondary load from the primary driving load, it is possible to freely construct the composite material of the load-bearing element, and in particular, it is possible to use many so-called shredded or short fibers rather than having the generally known continuous or endless fibers.

[0065] Because shredded or short fibers are randomly oriented within the matrix material of the composite material of the load-bearing element, knowledge about the orientation of mechanical stability and mechanical resistance is less predictable compared to the use of endless or continuous fibers. However, the construction of more complex structures (in particular complex circumferential and rotationally symmetric structures such as the load-bearing element of the present invention) can be produced much more easily than when using continuous fibers. In an aircraft, from the standpoint of stability prediction, the most important load is the drive load, which will be applied from the central rotor shaft through the hub to the multiple rotor blades. The main load or drive load can still be transmitted by using the enhanced stability provided by continuous fibers in the matrix material of the blade body. Here, additional mechanical stability (which is necessary and provided by the load-bearing element) is provided by a less expensive and less complex material using shredded fibers. Thereafter, the same lightweight advantage can be applied to the load-bearing element, but without limiting the complexity for producing the complex structure filled by that load-bearing element. In other words, the present invention enables a combination of the lightweight advantage of composite materials and the different complexities of different elements of the rotor structure. This combination is only possible due to the separation of primary load handling related to the driving load and secondary load handling related to the reaction load, which are at least partially isolated within the load bearing element.

[0066] When the load-bearing elements and rotor blades are formed as a monolithic structure, this can be an advantage of the present invention. While general combinations and force-transmission contacts can also be provided by mechanical contact or form-fit contact, etc., a material bridge of a monolithic structure between the load-bearing elements and the load-bearing sections of the blade body is an advantage. This can be achieved by a welding step, or by a melting step of thermoplastics, by using an adhesive to vulcanize the materials of both contacting elements. Monolithic combinations relate, among other things, to the use of thermoplastics as polymer matrix materials for the blade body and load-bearing elements. By melting the matrix materials together, a monolithic structure is realized, thereby enabling the bonding between the two elements with increased mechanical stability and also providing an easier production method, which will be described in more detail later.

[0067] Additionally, according to the present invention, the polymer matrix of the blade body and the polymer matrix of the load-bearing element can be selected from the same group of materials, in particular from thermoplastics. As already stated above, it is preferable to provide a monolithic structure of the blade body and the load-bearing element. One preferred method for realizing such a monolithic structure is the use of thermoplastics that can be welded or melted together. In particular, by using thermoplastics having similar or nearly identical melting points, the welding step and thereby the melting of both matrix materials of both elements is possible, and furthermore, a material connection between the load-bearing element and the blade body is achievable. These thermoplastics may include high-performance thermoplastics such as PEEK and PEKK. Alternatively, for similar or identical melting points, it may be advantageous if the melting point of the matrix material of the load-bearing element is significantly higher than the melting point of the matrix material of the blade body. In particular, when the production method described below is used, the second molding step includes injecting the molten high-temperature matrix material from the load-bearing element into a mold in which the already molded blade body is already installed. By injecting the matrix material of the load-bearing element at a high temperature, this material also provides heat transfer to the already hardened matrix material of the blade body in the load-bearing section. In other words, the injected high-temperature matrix material is capable of transferring such heat to the matrix material in the load-bearing section, at least partially, so that its temperature inherits the lower melting point of the matrix material in the load-bearing section, thereby enabling the material bridging for the monolithic structure to be realized in an easy and sufficient manner. In particular, no separate heat source is required to remelt the matrix material in the load-bearing section.

[0068] Additionally, a potential advantage of the rotor load bearing element of the present invention is that it includes a primary load bearing part that is in force-transmission contact with the load bearing section above the blade body, and a secondary load bearing part that is in force-transmission connection with the root section of the blade body. Improved mechanical stability is already achieved by a single primary load bearing part, but separation into two or more load bearing parts is also possible. This may be particularly advantageous if the central hub in the root section of the blade body uses the combination of the present invention of a monolithic structure between the load bearing element and the blade body. All load bearing parts include matrix material and shredded fibers for reinforcement. Furthermore, the driving load can be at least partially transmitted to the root section of the blade body via its secondary load bearing part.

[0069] According to the present invention, it is also advantageous if the load bearing element extends at least partially axially beyond the load bearing section of the blade body. This, in particular, allows the blade body to be enclosed within the load bearing section. Since the reaction load can be separated into, among other things, the centrifugal load, impact load, and air load acting on the blade body, at least the air load and impact load will result in a reaction load in the axial direction as well. Further extension of the load bearing element axially beyond the shape of the blade body allows for even better force transmission, receiving those reaction loads from the blade body and guiding them along the provided secondary load path. This extension is, among other things, on both sides of the blade body, i.e., in front of the blade body and behind the blade body following the air stream through the rotor.

[0070] Additionally, if the outer surface of the load-bearing element includes an aerodynamically functional shape, this can be an advantage. An aerodynamically functional shape must be understood as a shape with aerodynamic function. In other words, the outside of the load-bearing element is constructed to provide reduced aerodynamic drag and / or to provide guiding or even covering functionality for the airflow through the rotor. Since the engine's drive shaft in the duct is typically covered from the airflow, the load-bearing element can provide covering functionality as a secondary function. For example, the outer surface of the load-bearing element can provide a conical or curved shape to reduce airflow drag and provide the aforementioned guiding and / or covering functionality. Thus, functionality can be considered as passive aerodynamic functionality. Naturally, in addition to the relatively simple aerodynamic functionality discussed, more complex guiding functionality, such as providing additional rotor elements or guide fins, can also generally be implemented on the outer surface of the load-bearing element.

[0071] Further advantages may be realized if the rotor's load-bearing elements follow a rotationally curved shape that extends axially beyond the blade body, generating a pointed nose. Since the pointed nose can also be provided by a separate element, its integration into the load-bearing element further reduces complexity. Due to the fact that shredded reinforcing fibers can be used in the load-bearing element, the increase in complexity from integrating a conical pointed nose is unrelated to the complexity of the manufacturing method. In other words, the rotor's forward spinner can be integrated monolithically into the load-bearing element structure, thereby potentially significantly reducing the overall complexity for manufacturing and installation. In addition to the reduction in complexity, using the same lightweight material for the spinner as that which is part of the load-bearing element optimizes the advantages of the invention for lightweight structures with respect to the overall rotor.

[0072] Further advantages can be realized if the load-bearing element provides a hollow cavity. The hollow cavity is preferably located inside the load-bearing element and facing the root section of the blade body. Providing a hollow cavity means can be understood as having no material or less material in the hollow cavity, thereby further reducing the overall weight of the rotor. Additionally, the hollow cavity can be used to shield other mechanical elements (in particular the engine drive shaft) from the airflow entering the hollow cavity and interacting with those mechanical elements. The hollow cavity structure is combined, among other things, with the outer surface of the load-bearing element to provide an aerodynamically functional shape, as discussed above.

[0073] Further advantages may be realized if the blade body extends linearly or substantially linearly between the root section and the load-bearing section. This extension allows for a reduction in the complexity of the blade geometry in that section where interaction with airflow exists. In a preferred embodiment, the outer surface of the load-bearing element shields the airflow from entering the hollow cavity inside or near the root section, so that no airflow interaction occurs in that portion of the blade body between the load-bearing section and the root section. This allows for a reduction in structural complexity in that portion of the blade body, so that the blade body follows a substantially straight or linear extension. Furthermore, the structure of the blade body between the load-bearing section and the root section can be further reduced in terms of mechanical stability. The reason is that the reaction load resulting from the airflow and centrifugal load applied to the tip section of the blade body is at least partially transmitted to the secondary load part of the load bearing element, and less mechanical load needs to be guided through the portion of the blade body between the load bearing section and the root section. In other words, the blade body can be further reduced in terms of its weight with respect to the intermediate section of the body being discussed.

[0074] According to the present invention, a potential advantage is that the load-bearing element completely or at least substantially completely encloses the load-bearing section. In other words, the enclosing configuration can allow it to extend axially beyond the shape of the blade body, as discussed above. This allows for easier and better transmission of reaction loads into the secondary load path provided by the load-bearing section, and additionally, covers aerodynamic functions, as also discussed above. Substantially enclosing and covering the load-bearing section further allows for the form fit between the load-bearing element and the load-bearing section, as discussed below.

[0075] According to the present invention, a potential advantage is that the load-bearing section and the blade body include contact surfaces that extend at least partially along the axial direction of the rotor. This enables form-fitting as discussed above, thereby increasing and optimizing the transmission of centrifugal loads from the tip section of the blade body as reaction loads into the secondary load path provided by the load-bearing section. This can be used solely for form-fitting functionality or in combination with monolithic structures using material bonding between the matrix material of the load-bearing section and the load-bearing elements.

[0076] Additional advantages may be realized when the load-bearing element includes a rounded edge, at least in the contact section with the blade body. In particular, when a method including an injection molding step is used, the rounded edge can provide additional advantages to the injection molding process. Moreover, after production at the rounded edge, stresses generated during rotor use can be reduced, thereby avoided or at least reduced. The rounded edge according to the present invention includes, in particular, an edge radius between 1 mm and 10 mm.

[0077] According to the present invention, the rotor has the following geometric dimensions: - The radial extension of the load-bearing section compared to the overall radial extension of the blade body shall be between 20% and 50%. - The radial distance between the load-bearing section and the root section of the blade body must be between 20 millimeters and 50 millimeters. - The radial distance between the load-bearing section and the tip section of the blade body shall be between 100 mm and 150 mm. - The radial extension dimension of the blade body must be between 20 millimeters and 150 millimeters. - The number of rotor blades is between 3 and 35. If it includes at least one of these, that is also an advantage.

[0078] According to the present invention, it may be an additional advantage that all rotor blades are identical or substantially identical. In particular, this reduces the complexity and number of different parts, thereby reducing the cost of producing such rotors. This identity relates, among other things, to the material and form of the rotor blades.

[0079] A further object of the present invention is to provide a method for forming a rotor having the features of the rotor of the present invention. Such a method is - A step of forming a composite material comprising a polymer matrix and reinforcing continuous fibers into a cavity forming a rotor blade, wherein the rotor blade has a blade body, the blade body comprising a root section and a tip section at the opposite end of the blade body, the root section being connected to a drive shaft and for receiving the primary drive load, and - The steps of arranging multiple such rotor blades in a jig, - A step of forming a composite material comprising a polymer matrix and shredded reinforcing fibers into a cavity forming a load-bearing element, wherein such load-bearing element is formed in force-transmission contact with a plurality of blade body portions and in load-bearing sections located between a root section and a tip section. This includes providing a secondary load path for at least a portion of the reaction load on the blade body.

[0080] By forming the rotor of the present invention, the method of the present invention offers the same advantages as those discussed in detail with respect to the rotor of the present invention. As can be seen in the method of the present invention, using different fibers for different parts of the rotor allows for inexpensive and simple injection molding steps for producing separate elements of the rotor (in particular, the rotor blades and load-bearing elements). By using different fiber types for reinforcing fibers among the different elements of the rotor, different needs regarding load conditions can be addressed in a manner that still keeps the complexity and weight of the rotor low.

[0081] If an injection molding process is used for both molding steps, it can be advantageous. Among other things, this can involve the use of a polymer matrix material selected from thermoplastic material types. Among other things, this can be described as so-called overmolding, where, on the one hand, the injected fluid matrix material of the load-bearing elements remelts the matrix material of the rotor blades in their load-bearing sections, creating a material bond between the two parts.

[0082] According to the present invention, it may be an additional advantage if, during the second molding step, the molding temperature of the polymer matrix of the load-bearing element is set to be higher than the melting temperature of the polymer matrix material of the blade body. In other words, separate heating of the matrix material of the blade body is not required. This is because there is sufficient heat to be transferred into the molten matrix material of the load-bearing element, and this transferred heat can be used to remelt the polymer matrix material of the blade body in those load-bearing sections as it arrives in the load-bearing section in a molten form.

[0083] The present invention is further described in relation to the accompanying drawings, which schematically illustrate the following: [Brief explanation of the drawing]

[0084] [Figure 1] This figure shows an embodiment of the rotor of the present invention. [Figure 2] This figure shows some embodiments of the rotor of the present invention. [Figure 3] This figure shows a cross-section of an embodiment of the rotor of the present invention. [Figure 4] This figure shows a cross-section of a further embodiment of the rotor of the present invention. [Figure 5] This figure shows a further cross-section of a part of an embodiment of the rotor of the present invention. [Figure 6] This figure shows a further cross-section of a part of a further embodiment of the rotor of the present invention. [Figure 7] This figure shows the first step of the method of the present invention. [Figure 8] This figure shows further steps of the present invention. [Figure 9] This figure shows an aircraft equipped with the rotor of the present invention. [Figure 10] This figure shows a further embodiment of the rotor of the present invention. [Figure 11] Figure 10 shows the rotor of the present invention in a different diagram. [Figure 12] Figure 10 shows the rotor of the present invention in a different diagram. [Figure 13] Figure 10 shows a cross-sectional view of the rotor of the present invention. [Figure 14] Figure 10 shows the rotor of the present invention in a different diagram. [Figure 15] Figure 10 shows the rotor of the present invention in a different diagram. [Figure 16] This figure shows an embodiment of the rotor blade. [Figure 17] This figure shows the rotor blades of Figure 16 in a different diagram. [Figure 18] This figure shows an embodiment of the engine blade of the present invention. [Figure 19] This figure shows a cross-section along line AA from Figure 18, illustrating a different embodiment of the engine blade of the present invention. [Figure 20] This figure shows cross-sectional views of embodiments of the engine rotor, engine rotor assembly, and propulsion unit of the present invention. [Figure 21] This figure shows cross-sections of further embodiments of the engine rotor, engine rotor assembly, and propulsion unit of the present invention. [Figure 22] This figure shows the method steps of the present invention for manufacturing an engine blade. [Figure 23] This figure shows the method steps of the present invention for manufacturing an engine rotor assembly. [Modes for carrying out the invention]

[0085] Figure 1 shows one example of the rotor 10 of the present invention along the axial direction AD. This includes a plurality of rotor blades 20, which form the rotor 10 and thereby cause a fan engine 100 in a ducted structure. A drive shaft 110 is provided to supply a driving load to the rotor 10, and the drive shaft 110 is in driving force contact with the hub (in this case, all root sections 24 of the rotor blades 20). During operation, the drive shaft 110 rotates, thereby transmitting the driving load and causing all of the rotor blades 20 to rotate.

[0086] During operation, further load increases are applied to the blade body 22 (for example, impact loads, centrifugal loads CL, and / or air loads AL, as discussed later). Here, these secondary loads are at least partially absorbed by rotationally symmetric load bearing elements 30, which are connected to the entire blade body 22 in the load bearing section 28. In other words, here, at least a portion of the secondary loads are absorbed by the load bearing elements 30 and do not need to be transmitted to the root section 24 of the blade body 22.

[0087] Figure 2 shows details of the solution in Figure 1, focusing on one of the blade body sections 22. During operation, a secondary load in the form of a centrifugal load CL is applied to the tip section 26 of the blade body section 22 along the radial direction RD. To maintain the mechanism stably, a reaction load RL must be absorbed in opposition to this centrifugal load CL, where the reaction load RL is absorbed by a load bearing element 30 and can be transmitted through a load bearing section 28 in the circumferential secondary load path SLP. In other words, the reaction load RL is separated from the drive load DL (not shown in Figure 2), where it is removed from the mechanical stability of the root section 24 of the blade body section 22 of the rotor blade 20.

[0088] In Figure 3, a side section is shown. While rotating around a rotation axis extending along the axial direction AD, each rotor blade 20 receives an aerodynamic load AL resulting from the operation of the aircraft. Here, the aerodynamic load AL also generates a reaction load RL by a lever extending between the tip section 26 and the load-bearing section 28, and a further lever extending to the root section 24. In this embodiment, the load-bearing element 30 includes a primary load-bearing part 32 and a secondary load-bearing part 34, which can be connected to each other or separated from each other as shown in Figure 3. Thereafter, a hollow cavity 38 is created between the primary load-bearing part 32 and the secondary load-bearing part 34, reducing the amount of material required and, consequently, the overall weight of such a rotor 10. Additionally, as can be seen in Figure 3, the outer surface 31 of the load-bearing element 30 has an aerodynamically functional shape that guides the airflow and avoids air entering the hollow cavity 38. This can be seen, in particular, in correlation with the spinner, which can be a separate component or integrated into the load-bearing element 30, as shown in Figure 4.

[0089] Figure 4 shows the combination of the embodiment shown in Figure 3, integrated into the fan engine 100 according to Figure 1. Here, the spinner or pointed nose 36 is provided as a monolithic, integrated part of the load bearing element 30 (here, the primary load bearing part 32). It can also be seen that the drive shaft 110 is connected to the root section 24 or secondary load bearing part 34 of all the blade body sections 22 and applies the driving load to rotate the rotor 10.

[0090] In Figure 5, it can be seen that, along the indicated axial direction AD, the load-bearing section extends beyond the edges of the load-bearing section 28 at both ends in the axial direction AD. In particular, as can be seen in Figures 3 and 4, when bearing the reaction load RL resulting from the air load AL, the additional material is able to bear those reaction loads RL in a better and more stable manner.

[0091] Figure 6 shows the further increased complexity in the geometric shape of the load-bearing section 28. Here, in addition to the potential material bonding between thermoplastic matrix materials, this load-bearing section 28 includes a contact surface 29 that extends at least partially along the axial direction AD, so that a form fit along the radial direction RD can be achieved. This allows, in an even better embodiment, to integrate and transmit the centrifugal load CL to the reaction load RL of the secondary load path SLP of the load-bearing element 30.

[0092] Figures 7 and 8 illustrate a method according to the present invention, in which a single rotor blade 20 is formed in a blade mold 210. This is repeated multiple times to produce multiple identical or substantially identical rotor blades 20, which can then be placed in a fixture such as the rotor mold 220 shown in Figure 8. Here, molten material is injected by injection molding to produce load-bearing elements 30, remelting the load-bearing sections 28 of all contacted blade body portions 22. At the end of this injection molding process, it is possible to achieve not only form-fit functionality but also material bonding between the load-bearing sections 28 of the blade body portions 22 and the material of the load-bearing elements 30.

[0093] Figure 9 shows the aircraft designated by reference numeral 1. Aircraft 1 includes wings 2. In particular, aircraft 1 includes a left wing 2a and a right wing 2b, which extend along the transverse axis of the aircraft, on either side of the longitudinal axis of the fuselage 3. The transverse axis extends along the wingspan of wings 2a and 2b.

[0094] Multiple propulsion units 4 are mounted on both the left wing 2a and the right wing 2b. The propulsion units 4 are mounted on the rear portion of the wing 2 (particularly the rear end portion). The multiple propulsion units 4 are aligned along the transverse axis of the aircraft 1. The propulsion units 4 are mounted so as to be pivotable around the transverse axis. In particular, the propulsion units 4 are pivotable between a substantially horizontal angular position (cruising state) and a substantially vertical angular position (hovering state). Nine propulsion units 4 are provided on each of the wings 2a and 2b. Actuators are provided for each propulsion unit 4 to set the angular position of the propulsion unit 4 relative to the wing 2. In addition, the aircraft 1 includes canards 6. In particular, the left canard 6a and the right canard 6b are provided on both sides of the longitudinal axis in front of the wings 2a and 2b. A forward propulsion unit 5 is mounted on the rear portion of the canard 6 (particularly the rear end portion). Furthermore, the forward propulsion unit 5 is mounted on the canard 6 so as to be pivotable around an axis parallel to the transverse axis. Six propulsion units 4 are provided on the canards 6a and 6b, respectively. It is preferable that both the propulsion units 4 and the forward propulsion units 5 are of the ducted fan type. They can also be electrically driven by a rotating electromechanical device such as an electric motor.

[0095] The propulsion unit 4 and the forward propulsion unit 5 include flight control surfaces (flaps in this case) to generate additional lift. In other words, the outer casing of the propulsion unit 4 includes the lift-generating body. Each of the propulsion unit 4 and the forward propulsion unit 5 is capable of generating thrust by the rotation of a fan driven by an electric motor. It should be noted that each propulsion unit 4 and each forward propulsion unit 5 may have multiple ducted fans. In one embodiment of the present invention, at least one of the propulsion units 4, 5 includes a rotor 10 (hereinafter also referred to as engine rotor 10) having one or more rotor blades 20 (hereinafter also referred to as engine blades 20). Preferably, the rotor blades 20 can form the fan driven by an electric motor. Therein, at least one of the propulsion units 4, 5 is provided with an engine housing 120 with engine ducts, in which the rotor 10 is rotatably arranged around a rotation axis RA (see Figure 21 below). This allows, for example, the aforementioned configuration of a ducted fan to be derived.

[0096] Figures 10 to 17 illustrate further embodiments of the rotor 10 of the present invention in different diagrams. In this embodiment, the load-bearing element 30 includes a first load-bearing part 32 and a second load-bearing part 34. Additionally, Figures 16 and 17 show a single blade 20. Both figures show the same blade body in different diagrams, indicated by different square cross-sections. As can be seen in those figures, the load-bearing element 30 extends axially beyond the edge of the load-bearing section 28 of the blade body 22. Furthermore, a dovetail joint is part of this embodiment to provide a form fit to the shaft and to transmit the primary load to the rotor body.

[0097] To summarize some of the key aspects of the rotor 10 described above, the following provisions are provided.

[0098] 1. A rotor 10 for a ducted fan engine 100, wherein the rotor 10 is - A plurality of rotor blades 20, each having a blade body 22, the blade body 22 comprising a root section 24 and a tip section 26 at the opposite end of the blade body 22, the root section 24 being connected to the drive shaft 110 to receive the primary drive load and provide the primary load path, the plurality of rotor blades 20 being arranged circumferentially and forming the rotor 10, - To provide a secondary load path SLP for at least a portion of the reaction load RL of the blade body 22, a rotationally symmetric load bearing element 30 is in force-transmitting contact with a plurality of blade body 22 in a load bearing section 28 located between the root section 24 and the tip section 26. Includes, - The blade body portion 22 of the rotor blade 20 is made from a composite material including a polymer matrix and reinforcing continuous fibers embedded within the polymer matrix. - The rotor 10 is made of a composite material, the load-bearing element 30 of which includes a polymer matrix and shredded reinforcing fibers embedded within the polymer matrix.

[0099] 2. The rotor 10 according to Clause 1, characterized in that the load-bearing element 30 and the rotor blade 20 are formed as a monolithic structure.

[0100] 3. The rotor 10 according to any one of the preceding clauses, characterized in that the polymer matrix of the blade body 22 and the polymer matrix of the load bearing element 30 are selected from the same group of materials, in particular from thermoplastics.

[0101] 4. The rotor 10 according to any one of the preceding clauses, characterized in that the load bearing element 30 includes a primary load bearing part 32 that is in force transmission contact with the load bearing section 28 of the blade body 22, and a secondary load bearing part 34 that is connected to the root section 24 of the blade body 22 for force transmission.

[0102] 5. The rotor 10 according to any one of the preceding clauses, characterized in that the load-bearing element 30 extends at least partially in the axial direction AD beyond the load-bearing section 28 of the blade body 22.

[0103] 6. The rotor 10 according to any one of the preceding clauses, characterized in that the outer surface 31 of the load-bearing element 30 includes an aerodynamically functional shape.

[0104] 7. The rotor 10 according to Clause 6, characterized in that the load-bearing element 30 extends axially beyond the blade body portion 22 and follows a rotationally curved shape that generates a pointed nose 36.

[0105] 8. The rotor 10 according to any one of the preceding clauses, characterized in that the load-bearing element 30 provides a hollow cavity 38.

[0106] 9. The rotor 10 according to any one of the preceding clauses, characterized in that the blade body portion 22 extends in a linear or substantially linear manner between the root portion section 24 and the load-bearing section 28.

[0107] 10. The rotor 10 according to any one of the preceding clauses, characterized in that the load-bearing section 28 of the blade body 22 includes a contact surface 29 that extends at least partially along the axial direction AD of the rotor 10.

[0108] 11. The rotor 10 according to any one of the preceding clauses, characterized in that the load-bearing element 30 includes a rounded edge in at least the section that contacts the blade body 22.

[0109] 12. The rotor 10 has the following geometric dimensions: - The radial extension of the load-bearing section 28 is between 20% and 50% of the overall radial extension of the blade body 22. - The radial distance between the load-bearing section 28 and the root section 24 of the blade body 22 is between 20 mm and 50 mm. - The radial distance between the load-bearing section 28 and the tip section 26 of the blade body 22 is between 100 mm and 150 mm. - The radial extension dimension of the blade body portion 22 is between 20 mm and 1550 mm. - The number of rotor blades 20 is between 3 and 35. The rotor 10 according to any one of the preceding clauses, characterized by including at least one of the following.

[0110] 13. A method for forming a rotor 10 having the features described in any one of the clauses 1 to 12, the method being: - A step of molding a composite material comprising a polymer matrix and reinforcing continuous fibers into a cavity forming a rotor blade 20, wherein the rotor blade 20 has a blade body 22, the blade body 22 comprising a root section 24 and a tip section 26 at the opposite end of the blade body 22, the root section 24 being connected to a drive shaft 110 and for receiving a primary drive load DL, and - The steps of placing multiple such rotor blades 20 into the rotor mold 220, - A step of molding a composite material comprising a polymer matrix and shredded reinforcing fibers into a cavity forming a load-bearing element 30, wherein such load-bearing element 30 is formed in force-transmission contact with a plurality of blade body portions 22 in a load-bearing section 28 located between a root section 24 and a tip section 26, in order to provide a secondary load path SLP for at least a portion of the reaction load RL of the blade body portion 22. Methods that include...

[0111] 14. The method according to clause 13, characterized in that an injection molding process is used with respect to both molding steps.

[0112] 15. The method according to clause 13 or 14, characterized in that during the second molding step, the molding temperature of the polymer matrix of the load-bearing element 30 is set to be equal to or higher than the melting temperature of the polymer matrix material of the blade body 22.

[0113] Further aspects of the present invention are discussed below with reference to Figures 18 to 23. These figures illustrate different diagrams, aspects, and embodiments of the present invention. For example, references are made to the engine blade 20, which illustrate further aspects and embodiments of the rotor blade 20. Similarly, references are made to the engine rotor 10, which illustrate further aspects and embodiments of the rotor 10.

[0114] Figure 18 shows an optional configuration of an engine blade 20 for an aircraft engine 100. The engine blade 20 extends longitudinally along the blade axis BA. The engine blade 20 includes a blade body 22. The blade body 22 includes a blade surface 23, which, when in operation, preferably exerts an accelerating force on the air to be propelled. The blade surface 23 can extend circumferentially or chord-wise (for example, illustrated by line AA in Figure 18) between the transverse edges 22A, 22B. The transverse edges 22A, 22B may preferably mean, for example, the leading edge 22A and the trailing edge 22B of the engine blade 20 during operation. The blade surface 23 can extend longitudinally or spanwise between the tip section 26 and the root section 24.

[0115] Generally, the blade body 22 may preferably include two or more heating elements 40 (i.e., a first heating element 40A and a second heating element 40B) for heating different sections of the blade surface 23. As can be interpreted from Figure 18, the heating elements 40 can be positioned anywhere between the tip section 26 and the root section 24 of the blade body 22. The heating elements 40 can be arranged side by side circumferentially and / or longitudinally on the blade surface 23. The heating elements 40 may have the same configuration or different configurations. For example, the heating elements 40 may be of different sizes, as illustrated exemplary in Figure 18.

[0116] For example, the first heating element 40A can be positioned in the direction of the leading edge 22A. As illustrated by example, the first heating element 40A can extend spanwise along or parallel to the leading edge 22A between the tip section 26 and the root section 24. The first heating element 40A can be positioned proximal to the leading edge 22A. Preferably, the first heating element 40A can be configured and / or positioned so that it covers the chord-direction between 0% and 20% of the blade surface 23. Alternatively or additionally, the first heating element 40A can be positioned so that it covers the leading edge section of the blade surface 23 (which preferably extends uniformly from the leading edge 22A) in the chord-direction. Preferably, the leading edge section can be between 0% and 20% of the blade surface 23, or between 0% and 20% of the blade surface 23 on one side of the engine blade 20. One side can be the wing pressure side or wing pressure surface during operation.

[0117] Alternatively or additionally, the second heating element 40B may be positioned in the direction of the trailing edge 22B. As illustrated, the second heating element 40B may extend spanwise along or parallel to the trailing edge 22B between the tip section 26 and the root section 24. The second heating element 40B may be positioned proximal to the trailing edge 22B. Preferably, the second heating element 40B may be configured and / or positioned to cover the chord-direction between 20% and 100% of the blade surface 23. Alternatively or additionally, the second heating element 40B may be positioned to cover the trailing edge section of the blade surface 23 (which preferably extends uniformly from the trailing edge 22B) in the chord-direction. Preferably, the trailing edge section may be between 20% and 100% of the blade surface 23. Alternatively, the trailing edge section can be located between 20% and 100% of the blade surface 23 on one side of the engine blade 20.

[0118] Preferably providing the heating functionality of the engine blade 20 by two heating elements 40A, 40B facilitates a preferred heating control method for heating the engine blade 20. For example, the engine blade 20 can be heated in two stages. In the first stage, the first heating element 40A at least at the leading edge 22A, or only the first heating element 40A, can be operated. In the second stage, the second heating element 40B at least at the trailing edge 22B, or only the second heating element 40B, can be operated. For example, a control device of the aircraft 100 can control the heating elements 40A, 40B based on freezing conditions (e.g., external conditions (e.g., ambient temperature and / or humidity)) or other sensory indicators regarding ice formation on the engine blade 20. The control device can control the heating elements 40A, 40B so that the first stage switches to the second stage based on at least one of the freezing conditions exceeding an acceptable threshold.

[0119] This control method and configuration are advantageous because the heating function is activated first in the areas of the engine blade 20 that are prone to freezing due to direct collision with air, namely the leading edge 22A and the blade pressure side. Other areas are heated in a second stage only if necessary. Furthermore, this ensures the energy-efficient operation of the heating elements 40A and 40B. The proposed method is also advantageous because it ensures that the necessary properties regarding aeroelasticity and lightness are maintained while it leads to optimal de-icing and / or anti-freezing properties of the engine blade 20.

[0120] Alternatively or additionally, heating elements 40, 40A, and 40B may be provided and arranged in the spanwise direction (instead of the chordwise arrangement illustrated in Figure 18). Thus, the heating elements 40A and 40B can be arranged alternately in the longitudinal direction (for example, along the blade axis BA). Preferably, only the tip section 26 can remain uncovered by the heating elements 40, 40A, and 40B. Thereafter, areas of the blade surface 23 prone to ice formation (e.g., the blade pressure side and the leading edge 22A, etc.) can be provided with heating functionality.

[0121] Naturally, it is conceivable to implement such heating functionality with more than two heating elements 40, 40A, and 40B. Alternatively, a configuration with only a single heating element 40 is also conceivable. A single heating element 40 can cover a specific area of ​​the blade surface 23 and may include a controllable section to which energy is selectively supplied. Therefore, it is further conceivable that a heating control system can be formed by a heated blade 20 having one of these configurations and a control device configured to implement the aforementioned heating method.

[0122] The heating element 40 shown as an example may include heat transfer conduits 41. These may preferably be conductive and may extend along a curved extension path together with the blade surface 23. Preferably, a connection section 43 (e.g., an electrical connector for supplying power to the heating element 40) may be provided in the root section 24. A feed line 42 may connect the heat transfer conduits 41 to the connection section 43.

[0123] The heating element 40 is provided integrally with the blade body 22 within the blade surface 23. Therefore, the heating element 40 can be provided on top of the blade body 22 in an additive manufacturing method (e.g., 3D printing).

[0124] Figures 19A to 19C illustrate this embodiment of the engine blade 20. There, the blade body 22 is shown to include a layered structure 300, which includes a base layer 320 and a heating layer 340. The heating layer 340 includes a heating element 40, which is shown by a heat transfer conduit 41. The base layer 320 is provided for transmitting driving force from the drive shaft 110 through the root section 24 to the tip section 26.

[0125] As illustrated illustrative in Figures 19A to 19C, the base layer 320 may include several base support layers 321-324. These may be different components of the composite material (e.g., reinforcing fibers and polymer matrices).

[0126] As illustrated by Figures 19A to 19C, various configurations and arrangements of different layers of the layered structure 300 are possible.

[0127] For example, in Figure 19A, the heating layer 340 includes a carrier layer 341, and the heating element 40 is integrally provided on the carrier layer 341. As shown exemplary, the heat transfer conduit 41 can be printed on the carrier layer 341. The heating layer 340 is located on one of the base support layers 322-324 (i.e., base support layer 322). There, the heating layer 340 is integrally bonded to the base layer 320 via the carrier layer 341 by material bonding (e.g., through molding or curing). Furthermore, the heating element 40 is protected by a cover layer 330 that covers the heating layer 340 on the opposite side of the base support layer 322. The blade surface 23 is at least partially formed by the cover layer 330, and therefore the heating element 40 is provided beneath the blade surface 23.

[0128] Furthermore, in Figure 19B, the heating layer 340 also includes a carrier layer 341, and the heating element 40 is printed on the carrier layer 341. Unlike in Figure 19A, the carrier layer 341 is formed by the base support layer 322. The heating layer 340 is positioned between the other base support layers 321, 323, and 324. Thus, in this example, the heating layer 340 forms an integral part of the base layer 320, as it forms a layer of the base layer 320.

[0129] Figure 19C shows another example of an engine blade 20, where the heating element 40 (or more specifically, the heat transfer conduit 41) is printed directly onto the base layer 320. Thus, in this example, the heating layer 340 does not include a carrier layer. The heat transfer conduit 41 can be provided as a microstructure, and therefore, filler material (e.g., resin) may not be required to provide structural support between the heat transfer conduits 41.

[0130] Figure 20 illustrates a cross-section of the engine rotor 10 and cross-sections of propulsion units 4 and 5 equipped with the engine rotor 10. In this example, the engine rotor 10 includes an action factor supply device 51 for supplying power to a heating element 40 via a connection section 43. In particular, the action factor supply device 51 is provided as a slip ring 51. For example, a supply line 71 can be used to supply power to the slip ring 51 from a control unit 70. In the illustrated example, the slip ring 51 is integrally provided with the engine rotor 10. Furthermore, the slip ring 51 is positioned coaxially with the rotation axis RA and radially inward from the root section 24 with respect to the blade axis BA. The slip ring 51 is positioned to be movable relative to the engine housing 120 (see Figure 21) and preferably with respect to the drive shaft 110.

[0131] It is also conceivable to position the slip ring 51 in various other locations, such as laterally to the hub device 34 (previously referred to as the secondary load bearing part 34) relative to the rotation axis RA.

[0132] Figure 21 shows a cross-section of a further example of the engine rotor 10 and cross-sections of the propulsion units 4, 5. There, the acting force supply device 52 is provided as an inductive supply device 52 for wirelessly (preferably inductively) supplying electrical energy to the heating element 40. Energy transmission is illustrated exemplarily by arrows with dashed lines. As shown in Figure 21, the engine housing 120 includes the inductive supply device 52. The inductive supply device 52 can be located in the upstream or downstream section of the engine duct. Furthermore, the inductive supply device 52 can preferably be located upstream of the engine blade 20, however, a downstream configuration is also possible, as illustrated. In this example, the heating element 40 can correspondingly be provided as an electrical resonant circuit 45 for interacting with the inductive supply device 52.

[0133] Generally, the action factor supply devices 51 and 52 can supply or be configured to supply any type of energy suitable for the heating element 40 to de-ice the blade surface 23. For example, the action factor supply devices 51 and 52 can supply high-temperature gas or heat to the heating element 40. Furthermore, the action factor supply devices 51 and 52 can be connected to an energy source (e.g., a solar panel or a control unit 70).

[0134] Figures 22A to 22C illustrate different steps in manufacturing the engine blade 20. In Figure 22A, the heating element 40 is provided by printing heat transfer conduits 41 onto an uncured composite material layer as a carrier layer 341. For this purpose, an additive manufacturing device 400 (e.g., a print head) may be used. Furthermore, composite layers 321, 322, and 323 are provided and preferably cut to size for the engine blade 20. Naturally, other layers (e.g., a cover layer 330) may also be provided. Different composite layers 321, 322, 323, and 341 are laminated together as shown in Figure 22B. Figure 22C shows the joints of the different composite layers 321, 322, 323, and 341 forming an integrated engine blade 20.

[0135] Figures 23A to 23C illustrate different steps in a method for manufacturing an integral engine rotor assembly 90 (for example, as also illustrated in Figures 20 and 21). In Figure 23A, the hub device 34 for receiving and supporting the root section 24 of the engine blades 20 can be formed by compressing the respective materials. For example, a polymer matrix and shredded reinforcing fibers embedded within the polymer matrix can be used for this purpose. Figure 23B shows how multiple engine blades 20 are provided and positioned relative to the hub device 34. In addition, the slip rings 51 can be provided and positioned at this stage, or can already be provided and positioned in the step of forming the hub device 34. To form the integral engine rotor assembly 90, the engine blades 20, the hub device 34, and preferably the slip rings 51 are joined by injection molding or a composite press (exemplified by a press element 401). Processing in an autoclave is also possible.

[0136] The present invention is not limited to the embodiments described above, insofar as they are covered by the appended claims. All features of the embodiments described above can be combined in any possible way and can be provided interchangeably. [Explanation of Symbols]

[0137] 1...Aircraft, 2...Wing, 2a...Left wing, 2b...Right wing, 3...Fuselage, 4...Propulsion unit, 5...Propulsion unit, 6a...Left canard, 6b...Right canard, 10...Rotor, Engine rotor, 20...Rotor blade, Engine blade, 22...Blade body, 22A...Leading edge, 22B...Trailing edge, 23...Blade surface, 24...Root section, 26...Tip section, 28...Load bearing section, 29...Contact surface, 30...Load bearing element, 31...Outer surface, 32...Primary load bearing part, 34...Secondary load bearing part, Hub device, 36...Pointed nose, 38...Hollow cavity, 40...Heating element, 40A...First heating element, 40B...Second heating element, 41...Heat transfer conduit, 42…Feed line, 43…Connection section, 45…Electrical resonant circuit, 51…Slip ring, 52…Inductive feeding device, 70…Control unit, 71…Feed line, 90…Rotor engine assembly, 100…Fan engine, 110…Drive shaft, 210…Blade mold, 220…Rotor mold, 300…Layered structure, 320…Base layer, 321~324…Base support layer, 330…Cover layer, 340…Heating layer, 341…Carrier layer, 400…Additive manufacturing device, 401…Press element, CL…Centrifugal load, AL…Air load, RL…Reaction load, SLP…Secondary load path, RD…Radial direction, AD…Axial direction, BA…Blade axis, RA…Rotation axis

Claims

1. An engine blade (20) for an aircraft engine (100), wherein the engine blade (20) is Including the blade body (22), The blade body portion (22) is Blade surface (23) and A heating element (40, 40A, 40B) for heating at least a portion of the blade surface (23) and In the engine blade (20), including, The engine blade (20) is characterized in that the heating elements (40, 40A, 40B) are integrally provided with the blade body (22) within the blade surface (23).

2. The heating elements (40, 40A, 40B) include at least one heat transfer conduit (41), the at least one heat transfer conduit (41) is preferably conductive, and / or the at least one heat transfer conduit (41) extends along a curved extension path and / or together with the blade surface (23), and / or The engine blade (20) according to claim 1, characterized in that the heating elements (40, 40A, 40B) are provided on a portion of the blade body (22) in an additive manufacturing method, preferably in a 3D printing method or a deposition method.

3. The blade body portion (22) includes a layered structure, preferably a composite structure (300), and the composite structure (300) is A base layer (320) for transmitting driving force, A heating layer (340) including the aforementioned heating elements (40, 40A, 40B) Includes, Preferably, the heating layer (340) is disposed between the base layer (320) and a cover layer (330) that covers the heating layer (340) on the opposite side of the base layer (320), characterized in that the engine blade (20) according to claim 1 or 2.

4. The heating layer (340) is A carrier layer (341), preferably the carrier layer (341) is a foil or a film, and the carrier layer (341) The heating elements (40, 40A, 40B) are provided integrally on the carrier layer (341), preferably printed or deposited. Includes, The engine blade (20) according to claim 3, characterized in that the heating layer (340) is integrally bonded to the base layer (320) via the carrier layer (341) by material bonding.

5. The heating layer (340) forms an integral part of the base layer (320), Preferably, the base layer (320) includes at least two base support layers (321, 322, 323, 324), and the heating layer (340) is disposed between the base support layers (321, 322, 323, 324), or the heating layer (340) is disposed on one of the base support layers (321, 322, 323, 324). More preferably, the engine blade (20) according to claim 3 or 4, characterized in that the carrier layer (341) is one of the base support layers (321, 322, 323, 324).

6. The blade body portion (22) is made from a composite material comprising a polymer matrix and reinforcing continuous fibers embedded within the polymer matrix, and / or The engine blade (20) according to any one of the preceding claims, characterized in that the heating elements (40, 40A, 40B) are arranged between the tip section (26) and the root section (24) of the blade body (22).

7. The heating elements (40, 40A, 40B) include a connecting section (43) for supplying a heat generating factor, and the heating elements (40, 40A, 40B) extend between the tip section (26) of the blade body (22) and the root section (24) for attaching the engine blade (20) to the drive shaft (110), and the connecting section (43) is provided in the root section (24). Preferably, the engine blade (20) according to any one of the preceding claims, characterized in that the connecting section (43) is an electrical connector.

8. An engine rotor (10) for propelling an aircraft (1), An engine rotor (10) comprising at least one engine blade (20) as described in any one of the preceding claims.

9. A propulsion unit (4, 5) for propelling an aircraft (1), The engine rotor (10) according to claim 8, An engine housing (120) equipped with an engine duct, wherein the engine rotor (10) is rotatably disposed within the engine housing (120) around a rotation axis (RA) and A propulsion unit (4, 5) characterized by comprising:

10. A heat-generating factor supply device (51, 52) for supplying a heat-generating factor to the heating element (40, 40A, 40B) for heating the blade surface (23), preferably to the connecting section (43), The aforementioned factor supply devices (51, 52) are The aforementioned heat generation factor is an inductive supply device (52) for inductively supplying electrical energy to the heating elements (40, 40A, 40B), Preferably, the engine housing (120) includes the induction supply devices (51, 52), and the induction supply devices (51, 52) are In a part of the aforementioned engine duct, Upstream of the engine blade (20), It is arranged radially outward from the engine blade (20), The heating elements (40, 40A, 40B) are provided accordingly as an electric resonant circuit (45) for inductively receiving electrical energy from the inductive supply device (52), and / or The heat generation factor is a slip ring (51) for supplying electrical energy to the heating elements (40, 40A, 40B), The engine rotor (10) includes the slip ring (51), and preferably the slip ring (51) is integral with the engine rotor (10). The slip ring (51) is positioned to be movable relative to the engine housing (120), Preferably, the slip ring (51) is arranged coaxially with the rotation axis (RA) and preferably radially inward from the root section (24). The propulsion unit according to claim 9, characterized by comprising an action factor supply device (51, 52) including the above.

11. A method for manufacturing an engine blade (20) having a blade body (22) equipped with an integrated heating element (40, 40A, 40B) for heating at least a portion of the blade surface (23), Printing or depositing the heating elements (40, 40A, 40B) onto the blade body portion (22), A cover layer (330), preferably an adhesive, is applied to the heating elements (40, 40A, 40B) to form at least the portion of the blade surface (23) that is heated. A method characterized by including

12. A method for manufacturing an engine blade (20) having a blade body (22) equipped with an integrated heating element (40, 40A, 40B) for heating at least a portion of the blade surface (23), To provide composite layers (321, 322, 323, 324) for the engine blade (20), The heating elements (40, 40A, 40B) are provided and applied on at least one of the composite layers (340), The composite layers (321, 322, 323, 324, 340) for the engine blade (20) are laminated, The composite layers (321, 322, 323, 324, 340) are joined together to form the engine blade (20). A method characterized by including

13. The heating elements (40, 40A, 40B) are preferably provided and applied before stacking. Print the heating elements (40, 40A, 40B) on at least a portion of the composite layer. Printing at least one electrically conductive heat transfer conduit (41) directly onto the composite layer, The heating elements (40, 40A, 40B) are preferably placed on at least a portion of the composite layer together with the carrier layer (341) and cured, and / or Preferably, the heating elements (40, 40A, 40B) are arranged and bonded together with the carrier layer on at least a portion of the composite layer. Including, and / or, The manufacturing method according to claim 12, characterized in that the composite layer is an uncured layer.

14. The manufacturing method according to claim 12 or 13, characterized in that joining the composite layers (321, 322, 323, 324, 340) includes injection molding, composite pressing, curing, and / or bonding.

15. A method for manufacturing an integrated engine rotor assembly (90), wherein the method is: To provide a plurality of engine blades (20), wherein each engine blade (20) includes a blade body portion (22) having a blade surface (23), To provide a hub device (34) for receiving and supporting the root section (24) of the engine blade (20) In a method including, At least one of the engine blades (20) is manufactured by the method described in any one of claims 11 to 14, The engine blades (20) and the hub device (34) are arranged to form an engine rotor assembly (90), The engine blade (20) and the hub device (34) are joined together to form the aforementioned integrated engine rotor assembly (90). A method characterized by including

16. The manufacturing method according to claim 15, characterized in that joining the engine blade (20) and the hub device (34) includes injection molding, compound pressing, and / or joining.

17. Providing the hub device (34) includes forming the hub device (34), preferably the hub device (34) being made from a composite material comprising a polymer matrix and shredded reinforcing fibers embedded in the polymer matrix, by compression or injection molding. Preferably, the manufacturing method according to claim 15 or 16, characterized in that forming the hub device (34) includes providing and arranging action factor supply devices (51, 52) for supplying heat generation action factors to the heating elements (40, 40A, 40B) on the hub device (34), thereby providing the action factor supply devices (51, 52) that are integrated with the hub device (34).

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