Patent Application
20230348048
This application claims priority to European patent application No. EP 22170750.8 filed on Apr. 29, 2022, the disclosure of which is incorporated in its entirety by reference herein.
The present disclosure is related to a rotary wing aircraft with a fuselage and a landing gear.
Conventional landing gears of rotary wing aircrafts are either provided as retractable or partially retractable landing gears, or as fixed, i.e., non-retractable landing gears. Retractable or partially retractable landing gears are usually embodied as wheel-type landing gears, while fixed landing gears are usually embodied as skid-type landing gears.
Use of a wheel-type landing gear or a skid-type landing gear for a given rotary wing aircraft mostly depends on a respective size and weight of the given rotary wing aircraft. In other words, wheel-type landing gears are usually implemented with larger rotary wing aircrafts in order to allow for an improved ground handling of such larger rotary wing aircrafts, while light-to-medium rotary wing aircrafts are usually implemented with skid-type landing gears.
In general, respective landing gears of rotary wing aircrafts, in particular rotary wing aircrafts with vertical take-off and landing capabilities, are designed to allow absorption of vertical energy occurring during landing by means of elastic deformation during a regular or normal landing and plastic deformation during a hard or downward crash landing. The plastic deformation mainly contributes to energy dissipation of the vertical energy and shock absorption during such a hard or downward crash landing. For instance, such a plastic deformation during a hard or downward crash landing may reach values of up to 360 mm in vertical direction at the aft cross tube and of up to 460 mm at the forward cross tube in a helicopter-type rotary wing aircraft with a conventional skid-type landing gear having two skid tubes, a forward cross tube, and an aft cross tube. An illustrative skid-type landing gear with two skid tubes, a forward cross tube, and an aft cross tube is described in the document EP 2 610 170 A1.
Other landing gear designs that account for energy dissipation and shock absorption are also known. For instance, the document U.S. Pat. No. 3,144,223 describes a helicopter-type rotary wing aircraft with a fuselage and a non-retractable skid-type landing gear that comprises a pair of co-operable substantially identical shock absorbing portions which are secured to associated structural members provided at respective sides of the fuselage. Each shock absorbing portion comprises a torsion rod structure with two operatively distinct rod sections which extend in opposite directions toward outer ends from an intermediate reference location at which the torsion rod structure is operatively and non-rotatably secured by means of a bracket to the respectively associated structural member. The outer ends of the two operatively distinct rod sections are positioned in and supported by bearing means in a manner free to rotatably twist. The bearing means are also in the form of bearing brackets, respectively, which are secured to the respectively associated structural members. Furthermore, the two operatively distinct rod sections are welded or otherwise non-rotatably and fixedly connected to respective socket members adjacent to the outer ends and the socket members are rigidly mounted to laterally extending strut members, e.g., by means of threaded bolts.
The document U.S. Pat. No. 5,060,886, in turn, describes a non-retractable wheel-type landing gear for a vertical take-off and landing aircraft. The wheel-type landing gear includes a pair of spaced-apart fore and aft longitudinal tubes attached to a pair of short, spaced-apart cross tubes that are adapted to mount on structural saddles provided on an underside of the aircraft. The longitudinal tubes are connected together at the forward end for mounting a fully swiveling nose gear thereto. Aft main gear wheels are mounted on cantilever landing gear legs that extend laterally from the longitudinal tubes adjacent the aft cross tube. The cantilever landing gear legs are mounted by means of associated elastomeric bearings with cylindrical elastomeric springs to the longitudinal tubes. Landing loads on the wheel-type landing gear may be absorbed by resilient springs provided at the forward end, deflection of the gear legs and the aft cross tube, torsion of the fore and aft longitudinal tubes, or any combination of these actions. Resilient springs are used in the wheel-type landing gear to attenuate fatigue stresses in conjunction with the deflection of the longitudinal and cross tubes of the wheel-type landing gear.
The document U.S. Pat. No. 3,128,971 describes a non-retractable wheel-type landing gear that is mounted to a tip of a wing of an aircraft. The wheel-type landing gear includes a wheel that is rotatably mounted to one end of a suspension arm. The other end of the suspension arm is rigidly mounted to a cross shaft which is, thus, fixedly secured against rotation to the suspension arm. Both ends of the cross shaft are rotatably received and supported in associated bearings, which are respectively fixedly secured to front and rear walls of a spar that is arranged in the wing. The cross shaft is further fixedly connected to one end of a rectilinear torsion element in the form of an elongated tubular member which is rigidly mounted at its other end to the wing such that the other end is fixed against rotary motion. The elongated tubular member, however, may be subjected to a twisting action by an angular movement of the suspension arm.
However, in all of the above-described non-retractable landing gears the respective torsion elements are not separated from transmission of bending and normal loads. In other words, there is no functional separation in an underlying load path, since normal forces become bending with torsion and vice versa. Therefore, in the above-described non-retractable landing gears respectively absorbable energies are comparatively low and the landing gears as such only provide for a limited stiffness.
It is, therefore, an object of the present disclosure to provide a new landing gear and, more particularly, a new rotary wing aircraft with a landing gear that overcomes the drawbacks of the above-described landing gears.
This object is solved by a rotary wing aircraft with a fuselage and an at least partially non-retractable landing gear that comprises the features of claim 1. More specifically, according to the present disclosure the at least partially non-retractable landing gear comprises at least three individual support legs, wherein at least one individual support leg of the at least three individual support legs is mounted pivotally to the fuselage by means of an associated pivot bearing arrangement, wherein the at least one individual support leg is further connected to the fuselage via at least one torsion element with a predetermined elastic range, and wherein the at least one torsion element is adapted to act, in the predetermined elastic range, as a torsional spring, and, outside of the predetermined elastic range, as an energy absorber.
Advantageously, the at least partially non-retractable landing gear of the rotary wing aircraft according to the present disclosure fulfills the “no single point of failure” criteria of the EASA SC VTOL regulation due to provision of internal redundancies. In fact, every load path may simply be designed to be easily redundant.
Moreover, the at least partially non-retractable landing gear of the rotary wing aircraft according to the present disclosure advantageously consists of several major components forming a landing gear assembly. These major components can be arranged in various manners, whereby the share of functions is identical for each variant. Preferably, these major components comprise for the at least one individual support leg at least a leg box that forms the individual support leg as such, as well as a pivot bearing arrangement and a torsion element which connect the leg box to the fuselage of a given rotary wing aircraft, wherein the fuselage forms an aircraft interface. Optionally, the major components may further comprise a fairing.
More specifically, the leg box is preferably provided to transfer during landing all loads from ground to the fuselage (global forces in X, Y and Z direction). During normal landings, the leg box preferentially works in an associated linear elastic range as its main function is to transfer loads and not to absorb energy. The leg box may be made of metal and/or composite material. By way of example, the leg box may be mounted pivotally by means of the pivot bearing arrangement in respective lateral quarter shells of the rotary wing aircraft with respect to an axis which is at least approximately in parallel to a longitudinal axis of the rotary wing aircraft (X-direction) or, in the case of a tricycle-type landing gear, with an axis perpendicular to the longitudinal axis of the rotary wing aircraft. An associated pivot axis of the leg box is preferably arranged such that in case of a hard landing or a downward crash the leg box rotates around this pivot axis.
Advantageously, the leg box may be optimized with respect to different individual design points. By way of example, the leg box may be weight-optimized or stiffness-optimized without impacting other items of the at least partially non-retractable landing gear.
Preferably, the leg box acts individually and is connected to an associated solid base that may e.g., form a ground contact device. This ground contact device is preferentially provided to transmit loads from ground to the leg box and may be realized by any conventionally available device, such as e.g., a skid, a wheel, a snow shoe or even, in case of water or amphibious aircrafts, a floating device. In particular, use of a wheel enables e.g., rolling take-off and landings and further enables multiple applications of use. Preferentially, at least three punctual (e.g., wheels, snow shoes, etc.) or at least two linear (e.g., skids, etc.) ground contact devices are provided.
The pivot bearing arrangement is preferably provided to transfer during landing all forces in the vertical direction (forces in Z direction) from the leg box to the fuselage. More specifically, the main function of the pivot bearing arrangement is preferably to transfer shear forces from the leg box to the fuselage of the rotary wing aircraft.
The torsion element is preferably embodied as a bi-linear absorber with an inherent nonlinear force characteristic. More specifically, the torsion element preferentially works in two different stages or ranges respectively: In a regular or normal landing the torsion element preferably merely acts to introduce loads from the leg box into the fuselage in an associated elastic regime. Thus, the torsion element prevents the leg box from rotating around the pivot axis and functions as a torsion spring. In other words, in a normal landing the torsion element preferably stores occurring landing energy in a manner that is similar to an elastic spring effect. In a hard landing or downward crash, when a respectively occurring landing energy exceeds a predefined energy threshold or downward speed threshold, such as e.g., 1 m/s, the torsion element preferably dissipates respectively occurring kinetic energy by plasticization.
By way of example, the bi-linear absorber that forms the torsion element may be designed as one part as a simple torque loaded tube, i.e., as a torque tube made of a material with plasticization capabilities after exceeding yield strength in torsion or shear respectively, which implies that a material with a predetermined plasticization capability is used. This provides a possibility to tailor an underlying force deflection curve of the at least partially non-retractable landing gear almost as desired and to increase its absorbing capability up to crashworthiness.
However, in case the material used for the torque tube does not offer sufficient plasticization capability, a bolted connection that is adapted to dissipate energy by exceeding bearing strength may likewise be used. Alternatively, specific machine elements with a similar nonlinear force characteristic may also be used instead of a torque tube.
Advantageously, use of separate components for realizing the torsion element and the pivot bearing arrangement enables decoupling of the functions of energy absorption and load transfer such that the principles of lightweight construction may at least be applied to optimize design of the leg box. For instance, composite materials may be used or higher inertias due to increased support leg height. Furthermore, the torsion element as energy absorbing part may also be optimized as it has no load transfer duty. As a result, the at least partially non-retractable landing gear may be realized with a reduced overall weight.
Moreover, the use of separate components for realizing the torsion element and the pivot bearing arrangement enables a respective drag landing load transfer path from the energy absorption, hence, there is no negative impact on an underlying absorbing capability due to load superposition. This advantage mainly affects separation of a need for strength in longitudinal direction and a need for softness/compliance in vertical direction.
The fairing is preferably provided to embody the leg box with an aerodynamic shape in order to reduce aerodynamic drag. The fairing may be a separate component that is mounted to the leg box, or an integral part of the leg box. Preferentially, the fairing does not considerably contribute to stiffness and load carrying capacity of the leg box. However, as the fairing enables an overall design of the leg box which may at least approximate an airfoil, the overall aerodynamic drag of the at least partially non-retractable landing gear may be reduced significantly, and even more if it is designed without skids.
The aircraft interface is preferably provided to transmit forces from the leg box and respective moments from the torsion element to the fuselage. Accordingly, the aircraft interface must provide sufficient strength and stiffness in order to sustain all loads transferred via the leg box and the torsion element.
According to some aspects, the associated pivot bearing arrangement is provided to enable rotation of the at least one individual support leg about an associated rotation axis defined by the associated pivot bearing arrangement and to transfer shear forces from the at least one individual support leg to the fuselage, wherein the at least one torsion element is embodied with an inherent nonlinear force characteristic.
According to some aspects, the associated pivot bearing arrangement comprises at least one shear pin which is provided to transfer shear forces from the at least one individual support leg to the fuselage, wherein the at least one torsion element rigidly attaches the at least one individual support leg to the fuselage to limit at least in the predetermined elastic range of the at least one torsion element rotation of the at least one individual support leg about the rotation axis defined by the at least one shear pin.
According to some aspects, the at least one torsion element comprises at least one torque tube that rigidly attaches the at least one individual support leg to the fuselage to limit at least in the predetermined elastic range of the at least one torsion element rotation of the at least one individual support leg about the associated pivot bearing arrangement, wherein the at least one torque tube comprises in its length direction at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a constant outer diameter, a varying outer diameter, a constant wall thickness, a varying wall thickness, one or more holes, one or more slots, one or more grooves, one or more rims, one or more weak points, or one or more steps.
According to some aspects, the at least one torque tube is provided in its length direction with end flanges, wherein each end flange comprises in the length direction of the at least one torque tube at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a plurality of circular holes, a plurality of polygonal holes, or a plurality of slots.
According to some aspects, the at least one individual support leg comprises at least a first mounting flange and a second mounting flange, wherein the first mounting flange and the second mounting flange are pivotally mounted to the fuselage by means of the at least one shear pin.
According to some aspects, the at least one shear pin comprises a first shear pin for pivotally mounting the first mounting flange to the fuselage, and a second shear pin for pivotally mounting the second mounting flange to the fuselage.
According to some aspects, the first shear pin and the second shear pin are rigidly mounted to the fuselage.
According to some aspects, the at least one torque tube is rigidly mounted to the first mounting flange and to the fuselage.
According to some aspects, the at least one torsion element comprises an additional torque tube that rigidly attaches the at least one individual support leg to the fuselage to limit at least in the predetermined elastic range of the at least one torsion element rotation of the at least one individual support leg about the associated rotation axis defined by the at least one shear pin, wherein the additional torque tube is rigidly mounted to the second mounting flange and to the fuselage.
According to some aspects, the at least one individual support leg comprises a third mounting flange that is arranged along the associated rotation axis of the at least one individual support leg between the first mounting flange and the second mounting flange, wherein the at least one torque tube is rigidly mounted to the third mounting flange and to the fuselage close to the first mounting flange.
According to some aspects, the at least one torsion element comprises an additional torque tube that rigidly attaches the at least one individual support leg to the fuselage to limit at least in the predetermined elastic range of the at least one torsion element rotation of the at least one individual support leg about the associated rotation axis defined by the at least one shear pin, wherein the additional torque tube is rigidly mounted to the third mounting flange and to the fuselage close to the second mounting flange.
According to some aspects, the at least one shear pin connects the first mounting flange to the second mounting flange.
According to some aspects, the additional torque tube comprises in its length direction at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a constant outer diameter, a varying outer diameter, a constant wall thickness, a varying wall thickness, one or more holes, one or more grooves, one or more rims, one or more weak points, or one or more steps.
According to some aspects, the additional torque tube is provided in its length direction with end flanges, wherein each end flange comprises in the length direction of the additional torque tube at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a plurality of circular holes, a plurality of polygonal holes, or a plurality of slots.
It should be noted that the various different aspects of the present disclosure which are described above should not be understood as separate, isolated embodiments. Instead, any feasible combinations of the different aspects are likewise contemplated.
By way of example, in an illustrative realization of the inventive rotary wing aircraft with a fuselage and an at least partially non-retractable landing gear, the at least partially non-retractable landing gear may comprise at least three individual support legs, wherein each individual support leg is mounted pivotally to the fuselage by means of at least one shear pin which is provided to transfer shear forces from the individual support leg to the fuselage, wherein each individual support leg is further connected to the fuselage via at least one torque tube with a predetermined elastic range, the at least one torque tube being provided to transfer bending moments from the individual support leg to the fuselage, and wherein the at least one torque tube is adapted to act, in the predetermined elastic range, as a torsional spring, and, outside of the predetermined elastic range, as an energy absorber.
Illustrative embodiments of the disclosure are outlined by way of example in the following description with reference to the attached drawings. In these attached drawings, identical or identically functioning components and elements are labelled with identical reference numbers and characters and are, consequently, only described once in the following description.
Illustratively, the helicopter 1 comprises a fuselage 2 that forms, by way of example, a cabin 2a and a cockpit 2b. The fuselage 2 is connected to a tail boom 3 and comprises a lower fuselage region 2c, as well as an upper fuselage region 2d. Furthermore, the fuselage 2 is provided with an underside shell 2e, a portside shell 2f, and a starboard side shell 2g. Moreover, lateral quarter shells 4 are preferably provided in an area of the portside shell 2f and the starboard side shell 2g which is close to the underside shell 2e.
According to one aspect, the helicopter 1 comprises a landing gear 10. At this point it should be noted that only a part of the helicopter 1 with the fuselage 2 and the landing gear 10 is shown to illustrate the landing gear 10 and connection of the landing gear 10 to the fuselage 2 in more detail, while illustration of other well-known components is omitted, for simplicity and clarity of the drawing. For instance, neither the tail boom 3 is illustrated in greater detail for showing e.g., a suitable counter-torque device, nor a main rotor is shown, and so on.
Preferably, the landing gear 10 is at least partially non-retractable and, by way of example, the landing gear 10 is non-retractable. According to the present disclosure, the landing gear 10 comprises at least three individual support legs, wherein at least one individual support leg of the at least three individual support legs is mounted pivotally to the fuselage 2, as described in more detail below.
By way of example, and not for restricting the present disclosure accordingly, the landing gear 10 comprises four individual support legs, from which only two individual support legs are visible in
Illustratively, the individual support leg 10a comprises a leg box 11a that may be aerodynamically shaped. The leg box 11a may be connected via a support rod 12a to a solid base 13a that forms a ground contact device. By way of example, the solid base 13a is formed as a base plate or pad. Similarly, the individual support leg 10b comprises a leg box 11b that may also be aerodynamically shaped and that may be connected via a support rod 12b to a solid base 13b that forms a ground contact device. By way of example, the solid base 13b is also formed as a base plate or pad.
It should be noted that the individual support legs 10a, 10b are illustratively configured identically, at least within predetermined manufacturing tolerances. However, they may also distinguish from each other and be arranged differently on the fuselage 2. For instance, instead of providing the helicopter 1 with four individual support legs as illustrated, only three individual support legs may be provided. In this case, two of the three individual support legs may be mounted similar to what is shown in
Illustratively, the fuselage frames 40 include two separate individual webs 41, 42 which are respectively provided with an associated mounting plate 41a, 42a. The mounting plates 41a, 42a are preferably adapted to support the support leg 30, as described hereinafter.
The support leg 30 is preferably adapted for being mounted pivotally to the fuselage, i.e., the fuselage frames 40, by means of an associated pivot bearing arrangement 35. The pivot bearing arrangement 35 is illustratively provided to enable rotation of the support leg 30 about an associated rotation axis (44 in
The individual support leg 30 is preferably further adapted for being connected to the fuselage, i.e., the fuselage frames 40, via at least one torsion element 36 with a predetermined elastic range, wherein the at least one torsion element 36 is adapted to act, in the predetermined elastic range, as a torsional spring, and, outside of the predetermined elastic range, as an energy absorber. The at least one torsion element 36 is preferentially embodied with an inherent nonlinear force characteristic. Preferably, the at least one torsion element 36 rigidly attaches the support leg 30 to the fuselage, i.e., the fuselage frames 40, to limit at least in the predetermined elastic range of the at least one torsion element 36 rotation of the support leg 30 about the rotation axis (44 in
By way of example, the pivot bearing arrangement 35 comprises a single shear pin 35a and the at least one torsion element 36 is formed by two torque tubes 37, 38, which may be made of different materials. Each one of the torque tubes 37, 38 is preferably provided in its length direction, which is illustrated for both torque tubes 37, 38 with a dashed line 39, with associated end flanges 37a, 37b, 38a, 38b. Each end flange 37a, 37b, 38a, 38b may comprise in the length direction 39 at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a plurality of circular holes, a plurality of polygonal holes, or a plurality of slots, etc. Furthermore, each one of the torque tubes 37, 38 may be formed in its length direction 39 with at least one of a circular cross section, an elliptic cross section, a polygonal cross section, a constant outer diameter, a varying outer diameter, a constant wall thickness, a varying wall thickness, one or more holes, one or more slots, one or more grooves, one or more rims, one or more weak points, or one or more steps.
Preferably, the shear pin 35a is adapted for pivotally mounting a first mounting flange or lug 34a that is rigidly mounted to the torque tube 37, and a second mounting flange or lug 34c that is rigidly mounted to the torque tube 38, to the fuselage, i.e., the fuselage frames 40. The mounting lugs 34a, 34c are preferably integrated into a leg box 31 of the support leg 30, which e.g., implements one of the leg boxes 11a, 11b of
More specifically, the mounting lugs 34a, 34c are preferably formed as longitudinal end flanges of associated longitudinal frames or spars 31b, 31a of the leg box 31. Illustratively, the longitudinal frames 31a, 31b delimit the leg box 31 laterally and form together with upper and lower skins 32a, 32b which extend between the longitudinal frames 31a, 31b a box-shaped structure 32. Close to the longitudinal frame 31a a leading-edge fairing 33a may be mounted to the leg box 31 and close to the longitudinal frame 31b a trailing-edge fairing 33b may be mounted to the leg box 31 such that the latter is preferably aerodynamically shaped. However, instead of providing the leg box 31 with the fairings 33a, 33b, the leg box 31 as such may be formed with an aerodynamic shape.
Illustratively, the leg box 31 further comprises a transversal frame 31d that interconnects the longitudinal frames 31a, 31b close to the mounting lugs 34c, 34a. The transversal frame 31d is preferably interrupted by another longitudinal frame 31c that forms an associated mounting flange or lug 34b and that is arranged between the longitudinal frames 31a, 31b. Alternatively, the mounting lug 34b may be an integral part of the transversal frame 31d.
In an illustrative mounting procedure for mounting the support leg 30 and, more particularly, the leg box 31 to the fuselage, i.e., the fuselage frames 40, initially the end flange 37a of the torque tube 37 is rigidly mounted in a torque proof manner to the mounting plate 41a of the web 41, e.g., by means of a plurality of fasteners (51 in
Preferably, the shear pin 35a enables rotation of the support leg 30 about the shear pin 35a, i.e., about a respective rotation axis 44 that corresponds illustratively to the length direction 39 of the torque tubes 37, 38. However, the support leg 30 is connected via the torque tubes 37, 38 to the fuselage 2 such that the torque tubes 37, 38 allow rotation of the support leg 30 about the rotation axis 44, as illustrated with an arrow 45, in their associated predetermined elastic range by acting as torsional springs, i.e., by means of elastic deformation. Nevertheless, outside of the predetermined elastic range the torque tubes 37, 38 act as energy absorbers, i.e., by means of plastic deformation.
By way of example, the torque tubes 37, 38 are formed identically, at least within predetermined manufacturing tolerances. However, the torque tubes 37, 38 may also be provided with differing shapes, as described hereinafter.
According to the first variant, the torque tubes 37, 38 may have different tube lengths 61a, 61b. Alternatively, or in addition, they may have different tube diameters 62a, 62b.
According to the second variant, the torque tubes 37, 38 may still have different tube lengths 61a, 61b and/or different tube diameters 62a, 62b. However, alternatively, or in addition, they may have different tube wall thicknesses 63a, 63b.
It should be noted that the torque tube 38 of
The support leg 70 is embodied similar to the support leg 30 of
The support leg 70 is preferably adapted for being mounted pivotally to the fuselage, i.e., the fuselage frames 40, by means of two individual pivot bearing units 74a, 74b. The pivot bearing units 74a, 74b are illustratively provided to enable rotation of the support leg 70 about an associated rotation axis (44 in
By way of example, each one of the pivot bearing units 74a, 74b comprises a shear pin 75a, 75b that is illustratively integrated into an associated shear pin attachment 76a, 76b and provided to transfer shear forces from the support leg 70 to the fuselage, i.e., the fuselage frames 40. Preferably, the shear pin 75b is adapted for pivotally mounting the mounting flange or lug 34a that is rigidly mounted to the torque tube 37 to the fuselage, i.e., the fuselage frames 40, and the shear pin 75a is adapted for pivotally mounting the mounting flange or lug 34c that is rigidly mounted to the torque tube 38, to the fuselage, i.e., the fuselage frames 40.
In contrast to
In an illustrative mounting procedure for mounting the support leg 70 and, more particularly, the leg box 71 to the fuselage, i.e., the fuselage frames 40, initially the end flange 37a of the torque tube 37 is rigidly mounted in a torque proof manner to the mounting lug 34a of the leg box 71, e.g., by means of a plurality of fasteners (51 in
Preferably, the shear pins 75a, 75b enable rotation of the support leg 70 about the shear pins 75a, 75b, i.e., about the rotation axis 44. However, the support leg 70 is connected via the torque tubes 37, 38 to the fuselage 2 such that the torque tubes 37, 38 allow rotation of the support leg 70 about the rotation axis 44, as illustrated with the arrow 45, in their associated predetermined elastic range by acting as torsional springs, i.e., by means of elastic deformation. Nevertheless, outside of the predetermined elastic range the torque tubes 37, 38 act as energy absorbers, i.e., by means of plastic deformation.
The support leg 100 is embodied similar to the support leg 70 of
In an illustrative mounting procedure for mounting the support leg 100 and, more particularly, the leg box 71 to the fuselage, i.e., the fuselage frames 40, initially the end flange 37a of the torque tube 37 is rigidly mounted in a torque proof manner to the mounting lug 34a of the leg box 71, e.g., by means of a plurality of fasteners (51 in
Preferably, the shear pins 75a, 75b enable rotation of the support leg 100 about the shear pins 75a, 75b, i.e., about the rotation axis 44. However, the support leg 100 is connected via the torque tube 37 to the fuselage 2 such that the torque tube 37 allows rotation of the support leg 100 about the rotation axis 44, as illustrated with the arrow 45, in its associated predetermined elastic range by acting as a torsional spring, i.e., by means of elastic deformation. Nevertheless, outside of the predetermined elastic range the torque tube 37 acts as an energy absorber, i.e., by means of plastic deformation.
More generally, the main task of a landing gear of a rotary wing aircraft is to take, store, and/or dissipate respective kinetic energy of the rotary wing aircraft during landing. In general, the best practice to do so is a non-linear working curve, i.e., force deflection curve of a given landing gear assembly. This force deflection curve is usually created by various means, e.g., with a hydro/pneumatic strut (Oleo Damper) or due to plasticization of materials (for conventional skid-type landing gears).
Preferably, the force deflection curve is at least approximately bi-linear. Such a bi-linear behavior limits occurring forces and loads to an acceptable level, which is beneficial for a respective static sizing of the rotary wing aircraft's fuselage, and makes use of the stroke below the fuselage which is usually sufficiently available due to a comparatively big ground clearance below the fuselage. According to the present disclosure, this is achieved by using a material with high plasticization capabilities, such as e.g., aluminum, for realization of the at least one torsion element, i.e., the torque tube(s). After reaching yield strength in a hard or downward crash landing, the material is locally in the state of plastic deformation. As described above with reference to
More particularly, the force deflection curve 134 is illustratively composed of two essentially linear parts, i.e., linear approximations 138, 139. The linear approximation 138 illustrates an elastic domain 131 of the at least one torsion element of the at least partially non-retractable landing gear 10 of the rotary wing aircraft 1 of
It should be noted that modifications to the above described embodiments are within the common knowledge of the person skilled in the art and, thus, also considered as being part of the present disclosure. More particularly, the at least partially non-retractable landing gear 10 of the rotary wing aircraft 1 of
Moreover, a dedicated “weak point” may be predefined in the overall structure of the landing gear at the location of the torque tube(s). This location triggers the global behavior of the landing gear and may be taken as reference point for the design of the landing gear: all loads exceeding a certain level may be cut-off if exceeding a predetermined threshold value.
During normal operation, preferably the entire, at least partially non-retractable landing gear according to the present disclosure acts in its elastic regime, but as during operation local short-time overload may take place, the torque tube(s) takes these overload forces and dissipates small fractions of energy by plasticization. Accordingly, during normal operation the plasticization of the torque tube(s) may be used as an indicator and measurement device for the loads which occurred during operation. As all other parts are preferably still operating in their elastic range when the torque tube(s) already operates in its plastic range, only the torque tube(s) has to be exchanged from time to time.
The shown innovative arrangement is further beneficial, in that, should a hard or downward crash landing within the design envelope occur, the deformed element, i.e., the torque tube(s) is of simple and easily exchangeable nature and, thus, may be exchanged quickly. Further benefit is achieved as the torque tube(s), by magnitude of its deformation, provides a clear indication of forces having acted on the inventive landing gear and, thus, allows simple classification of incident severity and associated required return-to-service actions.
In addition, the inventive landing gear provides a solution to provide multiple landing gear configurations (wheeled, skid, etc.) on a common airframe structure without requiring significant structural changes in manufacturing (cost) or inclusions of structural provisions (weight) for multiple configurations.
| Number | Date | Country | Kind |
|---|---|---|---|
| 22170750.8 | Apr 2022 | EP | regional |
