Driverless hybrid inflatable aircraft

Abstract

A hybrid unmanned aerial vehicle (1) is described, configured for the combined and optimized use of aerostatic and aerodynamic forces, comprising an inflatable body (10) comprising an outer casing (11) and a load-bearing structure (20) inside said outer casing (11), said inflatable body (10) being adapted to adopt a closed-wing operating configuration.

Description

Technical Field

[0001] This invention relates to the technical field of unmanned aerial vehicles. Background Technology

[0002] Unmanned aerial vehicles are known. For example, in recent years there has been growing interest in tropospheric and stratospheric unmanned aerial vehicles, particularly those in stratospheric platforms also known as HAPS (High Altitude Pseudo-Satellites), without introducing any restrictions. This interest is driven by analysis of potential applications that would benefit from the advantages of using unmanned aerial vehicles, such as monitoring land security, precision agriculture, telecommunications, and environmental monitoring.

[0003] Inflatable aircraft platforms in the troposphere and stratosphere have four main advantages over composite or metal platforms: reduced structural weight, fewer aeroelastic problems due to structural properties, the ability to use lighter gases than air as a retainer, and reduced transport volume when deflated.

[0004] Unlike satellites, the HAPS platform offers the possibility of continuous observation of the Earth's surface at a local scale, with virtually no revisit time, and the possibility of observing the Earth's surface in the vicinity, i.e., from an altitude significantly lower than that of satellites. This provides image resolution that is much higher than that provided by satellite sensing, and even requires less installation and has a lower cost-effectiveness.

[0005] HAPS is a platform that can typically operate at an altitude of about 18 km to 20 km above the Earth's surface (lower stratosphere), and HAPS can usually operate continuously for several months using photovoltaic energy.

[0006] The flight altitude (18 km to 20 km) is of great interest because it is well above the range of altitudes relevant to commercial air traffic. Therefore, the impact of the HAPS platform on airlines is limited to the climb to mission altitude and the descent to landing base altitude. Furthermore, from a meteorological perspective, statistical analysis of current wind conditions shows that winds are weakest within this altitude range. The stratosphere has a temperature distribution that increases with altitude (unlike the troposphere). This makes this part of the atmosphere stable, thus preventing the formation of updrafts and turbulence.

[0007] In recent years, several recommendations have been made regarding HAPS platforms. These recommendations can be categorized into three main platform types. The first type, also from a temporal perspective, is represented by stratospheric balloons. The second type is represented by fixed-wing platforms, and finally, the third type is represented by airship platforms.

[0008] Stratospheric balloon and airship platforms are based on balancing weight using only aerostatics (lighter than air), while fixed-wing platforms use only aerodynamics (higher than air) to balance weight.

[0009] The payload-to-total-weight ratio of balloon-type platforms is relatively high. On the other hand, these balloon-type platforms typically lack propulsion and other directional control systems, thus offering no possibility of setting a predetermined trajectory, which is therefore determined by wind conditions at different altitudes.

[0010] On the other hand, fixed-air platforms increase aerodynamic efficiency based on configurations with very large wing extensions, and are characterized by their very lightweight structure. Due to this very lightweight structure, the total weight of fixed-air platforms must be limited, thus reducing the usable payload weight to a few kilograms (5 kg to 25 kg). Furthermore, the very lightweight structure of these platforms can cause aeroelastic problems during low-altitude, ascent, and descent phases.

[0011] On the other hand, compared to fixed-wing platforms, airship platforms are designed for very high payloads (>200 kg) and have very large dimensions and very high maximum takeoff weights (MTOW>3000 kg to 5000 kg). This MTOW is mainly due to the presence of gasbags, which form a second inner shell almost the same size as the outer shell, necessary to compensate for gas expansion during altitude changes (the volume at one altitude can be up to 20 times the volume at sea level). The enormous size of stratospheric airships makes ground operation difficult, especially in adverse weather conditions, and requires specialized infrastructure for deployment and shelter, making the tactical use of such stratospheric airships virtually impossible.

[0012] Therefore, the analysis of the aforementioned existing stratospheric platforms reveals two main aspects: First, the operation of such platforms is heavily influenced by weather conditions, leading to the belief that they can only be deployed from a limited number of bases worldwide. Second, the usable weight for payload is limited; for fixed-wing platforms, this weight is restricted to over 25 kg, while for airship-type platforms, a payload exceeding 200 kg is truly practical. Summary of the Invention

[0013] The object of the present invention is to provide an alternative type of aircraft to the platform described above, characterized by allowing efficient absorption of torsional loads and allowing for reduced aerodynamic drag without a large wingspan aerostructure configuration.

[0014] According to an aspect of the invention, as an alternative to or supplement to the foregoing objectives, another objective is to provide an aircraft that, relative to the platform discussed above with reference to the prior art, allows for reductions in the size and / or volume and / or cost and / or operational disadvantages of the aircraft.

[0015] According to an aspect of the invention, as an alternative to or supplement to the above objectives, another objective is to provide an aircraft that allows for the resolution or at least partial overcoming of the disadvantages discussed above with reference to the prior art platforms.

[0016] According to an aspect of the invention, as an alternative to or supplement to the above objectives, another objective is to provide an aircraft that allows for a reduction in the weight, size, and cost of an aircraft for the same payload relative to the prior art, and thus such an aircraft can also be used for tactical operations.

[0017] According to an aspect of the invention, as an alternative to or supplement to the foregoing objective, another objective is to provide an aircraft that allows for improved system reliability relative to the systems of the prior art described above.

[0018] According to an aspect of the invention, as an alternative to or supplement to the above-mentioned objective, another objective is to provide an aircraft capable of carrying a payload in the range of 5 kg to 100 kg required for many applications.

[0019] These and other objectives are achieved by an unmanned hybrid inflatable aircraft as defined in this invention. Attached Figure Description

[0020] The invention will be better understood through the following detailed description of embodiments of the invention, taken in an exemplary and non-limiting manner with reference to the accompanying drawings, in which:

[0021] - Figure 1 A schematic perspective view of the front of a hybrid-powered inflatable aircraft according to a presently preferred embodiment is shown; the aircraft is shown without a payload.

[0022] - Figure 2 It shows Figure 1 A schematic perspective view of the front of the internal load support structure of the aircraft.

[0023] - Figure 3 It shows Figure 1 A plan view of the front of the aircraft; the payload is also shown in the diagram.

[0024] - Figure 4 It shows Figure 1 A plan view of the rear of the aircraft in the diagram; the aircraft is shown without a payload.

[0025] - Figure 5 It shows Figure 1 A side plan view of the aircraft; the aircraft in this figure is shown without a payload;

[0026] - Figure 6 It shows Figure 1 A top view of the aircraft in the figure; the aircraft in this figure is shown without a payload;

[0027] - Figure 7 It shows Figure 1 A plan view of the bottom of the aircraft; the aircraft in this diagram is shown without a payload.

[0028] - Figure 8 It shows Figure 1 A perspective view of a portion of the aircraft, in which part of the aircraft's outer shell has been removed;

[0029] - Figures 9A to 9C These are respectively in the initial takeoff configuration, intermediate configuration, and operational mission configuration. Figure 1 A schematic plan view of the front side of the aircraft. Detailed Implementation

[0030] Similar or equivalent elements in the accompanying drawings will be indicated by the same reference numerals.

[0031] It should also be noted that the terms "lower," "upper," "horizontal," and "vertical" used in the following description of the components of the aircraft according to the invention are intended to indicate the normal operating condition of such an aircraft at a specified mission altitude. Again, the terms "front" and "rear" used in the following description of the components of the aircraft according to the invention are intended to indicate the forward direction X1 of such an aircraft in normal operating condition at a specified mission altitude. Figure 6 Furthermore, the terms "radial outer" and "radial inner" used to describe a portion of the aircraft platform according to the invention are intended to indicate the axis of the opening 12, which will be described later (see, for example, [reference needed]). Figure 3 ), wherein the axis is centered and orthogonal to the opening 12.

[0032] First refer to Figures 1 to 7The unmanned hybrid-powered aircraft according to the currently preferred embodiment is generally indicated by reference numeral 1 in the accompanying drawings. Specifically, according to the embodiment, aircraft 1 is a stratospheric platform, and more preferably a HAPS (High Altitude Pseudo-Satellite) stratospheric platform. Furthermore, it should be noted that, for the purposes of this description, the term "hybrid-powered" used to define the aircraft indicates the fact that the aircraft is constructed for the optimized combination and balance of aerostatic and aerodynamic applications. The term "optimized" refers to selecting a buoyancy ratio (i.e., the percentage of aerostatic force to total force) to minimize the total weight and footprint of the platform, fix the payload, and achieve the minimum altitude achievable using only aerostatic force. Specifically, the hybrid-powered aircraft according to this specification is constructed to jointly utilize both aerostatic buoyancy and lift. In other words, it should also be noted that aircraft 1 is conveniently constructed to utilize aerodynamics not only to control aircraft 1 during each flight phase but also to balance the weight of the aircraft itself. According to the embodiment, aircraft 1 is a tactical aircraft, meaning that the aircraft can be easily transported in a standard container and deployed directly in operational scenarios. This is thanks to its small size and MTOW (Maximum Takeoff Weight). (See reference) Figure 5 and Figures 6 to 7 According to the implementation, specifically, the aircraft 1 has a maximum length L1 or wingspan L1 ranging from 8 m to 25 m, a maximum payload MTOW ranging from 30 kg to 400 kg, and a payload 60 ranging from 5 kg to 100 kg. Generally, the aircraft 1 is suitable for use in land surveillance. Specifically, the aircraft 1 can be used for a variety of different applications, such as, but not limited to, border surveillance, environmental monitoring, precision agriculture, telecommunications, homeland security, and emergency support.

[0033] According to embodiments, as will be understood in more detail below in this specification, the aircraft 1 has a combination of inflatable structural elements with different internal pressures, these inflatable structural elements being preferably connected to rigid substructures (e.g., rigid substructures made of composite materials, aluminum, etc.), such as cabins, engine mounts, movable control surfaces, etc. According to embodiments, all inflatable elements of the aircraft 1 described below in this specification comprise laminated material components consisting of a gas-retaining layer, a structural layer, and a protective layer.

[0034] Refer again Figures 1 to 7 The aircraft 1 includes an inflatable body 10. The inflatable body 10 includes an outer shell 11 or outer skin 11 and a load-bearing structure 20 within the outer shell 11. Conveniently, the inflatable body 10 is adapted to a closed-wing operating configuration. Figure 1 , Figures 3 to 6 , Figure 7 and Figure 9CSpecifically, in this closed-wing operating configuration, the inflatable body 10 has an annular shape extending around the through-hole 12. It should be noted that, according to the embodiment, the aforementioned closed-wing operating configuration corresponds to a design or mission configuration in which the platform 1 is filled with gas, preferably helium. According to the embodiment, in the closed-wing operating configuration, thanks to its specific aeronautical structural construction, the aircraft 1 can very effectively generate approximately 60% to 80% of the lift and 40% to 20% of the static buoyancy of the aircraft 1's total weight, which is filled with a gas lighter than air, preferably helium as described above.

[0035] According to an embodiment, in a closed-wing operating configuration, the inflatable body 10 includes a first arched portion 10A, which is adapted to define an arched wing 10A having a leading edge 102A and a trailing edge 101A. Preferably, the arched wing 10A is a semi-elliptical or substantially semi-elliptical wing. Furthermore, in the closed-wing operating configuration, the inflatable body 10 includes a second straight portion 10B, which is adapted to define a straight wing 10B having a leading edge 102B and a trailing edge 101B. Specifically, the arched wing 10A is the upper wing, and the straight wing 10B is the lower wing. In the closed-wing operating configuration, the inflatable body 10 also includes a third connecting portion 10C and a fourth connecting portion 10D disposed at opposite ends of the straight wing 10B. The connecting portions 10C and 10D are adapted to connect the arched wing 10A and the straight wing 10B to each other. In other words, the connecting portions 10C and 10D correspond to the opposite end portions or ends of each of the wings 10A and 10B. In fact, by means of the connecting portions 10C and 10D, the straight wing 10B structurally and aerodynamically closes the arched wing 10A.

[0036] Reference Figure 5 and Figure 7 According to one embodiment, in a plan view of the aircraft 1 in the above-described operating configuration from below, the leading edges 102A and 102B are aligned with each other, while the trailing edges 101A and 101B are offset from each other. However, according to an alternative embodiment, in a plan view of the aircraft 1 in the above-described operating configuration from below, the leading edges 102A and 102B can be offset from each other to improve the stability and controllability characteristics of the aircraft. In fact, by shifting the pressure center of the domed wing 10A and the straight wing 10B, longitudinal moments relative to the center of gravity can be eliminated, for example.

[0037] According to one embodiment, control surfaces 80A and 80B are associated with the arched wing 10A and the straight wing 10B, respectively. Control surfaces 80A and 80B perform control and trim functions for the aircraft 1. According to another embodiment, control surfaces 80A and 80B can also be rigid structures made of composite materials.

[0038] According to one embodiment, the aircraft 1 includes a solar panel 70, preferably a flexible solar panel 70, which may be associated with an arched wing 10A and / or a straight wing 10B. According to another embodiment, the aircraft 1 includes a pair of housings or nacelles for a payload 60 and preferably also for batteries and avionics systems, located at a third connection portion 10C and a fourth connection portion 10D.

[0039] According to an embodiment, the aircraft 1 includes at least one propulsion system 51, 52. The propulsion systems 51, 52 preferably include a pair of forward thrusters 51, spaced apart from each other and associated with an arched wing 10A. Furthermore, the propulsion systems 51, 52 include a pair of forward thrusters 52, spaced apart from each other and associated with a straight wing 10B. However, according to an embodiment, more than four thrusters 51, 52 may be present. According to an embodiment, the thrusters 51 are aligned with each other along the main extension direction of the arched wing 10A, while the thrusters 52 are aligned along the main extension direction of the straight wing 10B. For example, in… Figures 6 to 7 As can be seen, according to the embodiment, thrusters 51 are relatively closer together, while thruster 52 is relatively farther away than thruster 51. According to the embodiment, thrusters 51 and 52 include helical thrusters. More specifically, according to the embodiment, thrusters 51 and 52 include driving helical thrusters. Preferably, thrusters 51 and 52 include electric motors 51 and 52. Specifically, according to the embodiment, the propulsion systems 51 and 52 are all electric and the power generation is based on the aforementioned solar panel 70. More specifically, according to the embodiment, the aircraft 1 is designed to be completely energy self-sufficient and capable of staying in the air for several weeks. Preferably, the energy storage required for nighttime flight is based on high-energy-density batteries, such as lithium polymer batteries (Li-Po), lithium-ion batteries (Li-ION), lithium-sulfur batteries (Li-S), etc.

[0040] The specific closed-wing operating configuration of aircraft 1 has significant structural and aerodynamic advantages. Structurally, the arched wing 10A has a small portion of the load distributed on the arched wing (partially supported by the static thrust generated by the internal gas). For example, according to an embodiment, the load distributed on the arched wing 10A is due to the weight of the wing 10A itself and preferably due to the solar panel 70 and / or the thruster 51 and / or the control element 40 (described in more detail below). Figure 8The load is generated by the presence of the control element 40 shown below. However, most of the load will be concentrated at the end portion or tip of the wing 10A, i.e., at the connecting portions 10C, 10D, where the payload 60 and preferably the battery and avionics system are arranged. According to the embodiment, the wing 10A will bear the main tensile load due to the arched structure and specific load configuration in the connecting portions 10C, 10D, which is preferably matched to the material used to manufacture the wing 10A itself, preferably primarily fabric. According to the embodiment, the straight wing 10B will structurally have only the following distributed load: the weight of the wing 10B itself and preferably the weight of the propeller 52 and / or the corresponding control element 40 (which will be described below). The straight wing 10B will contribute only a portion of the load concentrated at the end portion or tip of the straight wing, i.e., at the connecting portions 10C, 10D. According to the embodiment, the straight wing 10B must primarily support the bending load, which is reduced by both the static buoyancy and aerodynamic load generated by the arched wing 10A at the corresponding end portions, i.e., at the connecting portions 10C and 10D. Again, from a structural perspective, the closed-wing operating configuration effectively addresses the need to absorb torsional loads, a typical feature of classic fixed wings with free ends. Furthermore, this load is preferably absorbed by the load-bearing structure 20 and the outer skin 11.

[0041] From an aerodynamic perspective, the closed-wing operational configuration represents the optimal solution for reducing induced drag without having to consider a wide wingspan. The aerodynamic load on the sculpted wing 10A is greater than that on the straight wing 10B. Furthermore, due to its sculpted configuration, the sculpted wing 10A generates lateral aerodynamic forces in addition to lift, which tractions the straight wing 10B and helps to support the bending load of the straight wing 10B.

[0042] See Figure 2 According to the embodiment, the load support structure 20 is a ring structure, and the load support structure extends through the arched wing 10A, the straight wing 10B, and the aforementioned third connecting part 10C and fourth connecting part 10D.

[0043] According to one embodiment, the load support structure 20 includes at least one annular main wing spars 201. Advantageously, according to one embodiment, the at least one annular main wing spars 201 is an expandable wing spars 201. In an exemplary embodiment, the load support structure 20 includes a single main wing spars 201. According to one embodiment, in a closed-wing operating configuration, the at least one main wing spars 201 includes an arched main wing spars portion 201A and a straight main wing spars portion 201B. The arched main wing spars portion 201A is associated with an arched wing 10A, while the straight main wing spars portion 201B is associated with a straight wing 10B. Specifically, the arched main wing spars portion 201A is a tubular portion, preferably having a circular cross-section, which tapers from the center of the portion 201A toward the third connecting portion 10C and the fourth connecting portion 10D. Furthermore, the straight main wing spars 201B is a tubular portion, preferably having a circular cross-section, which tapers from the center of the portion 201B toward the third connecting portion 10C and the fourth connecting portion 10D. Preferably, the cross-section of the portion 201B varies in particular with the percentage change in the thickness of the wing profile 10B.

[0044] Refer again Figure 2 According to one embodiment, the load-bearing structure 20 includes at least one annular auxiliary wing spars 202-204, which have a cross-section smaller than that of the main wing spars 201. According to another embodiment, the at least one auxiliary wing spars 202-204 are advantageously expandable wing spars. According to another embodiment, the at least one auxiliary wing spars 202-204 includes a rear edge wing spars 202, a front edge wing spars 204, and a middle wing spars 203. Specifically, the rear edge wing spars 202 are arranged at the rear edges 101A, 101B of the arched wing 10A and the straight wing 10B. The front edge wing spars 204 are arranged at the front edges 102A, 102B of the arched wing 10A and the straight wing 10B. The middle wing spars 203 are inserted between at least one main wing spars 201 and the rear edge wing spars 202. According to one embodiment, the intermediate spar 203 can be positioned at the point where the thickness percentage of wings 10A and 10B is greatest, and the main spar 201 is positioned at a distance from the trailing edge spar 202, this distance being approximately 65% ​​to 85% of the distance between the trailing edge spar 202 and the leading edge spar 204. Advantageously, according to one embodiment, all auxiliary spars 202-204 are inflatable spars. According to one embodiment, the at least one main spar 201 and the at least one auxiliary spar 202-204 are fluidly connected to each other and are preferably inflated to the same pressure.

[0045] Refer again Figure 2According to one embodiment, the inflatable body 10 includes at least one annular chamber 31, 31A, 31B, 32, defined by an outer skin 11 and extending into the arched wing 10A and the straight wing 10B. According to another embodiment, the inflatable body 10 specifically includes a plurality of annular chambers 31, 31A, 31B, 32. According to another embodiment, such a plurality of annular chambers 31, 31A, 31B, 32 include chambers in fluid communication with each other. For example, according to another embodiment, the outer skin 11 is fixed only, preferably glued to the leading edge spar 204 and the trailing edge spar 202, thereby allowing gas to pass between the annular chambers 31, 31A, 31B, 32. According to an alternative embodiment, if there are ailerons, for example, provided with corresponding chambers, the outer skin 11 may also be fixed, preferably glued to the intermediate spar 203. According to an embodiment, in a closed-wing operating configuration, the at least one annular chamber 31, 31A, 31B, 32 is inflated to a pressure lower than the pressure at which the at least one main spar 201 is inflated, and, if provided, the pressure at which the at least one auxiliary spar 202-204 is inflated. According to a convenient embodiment, the at least one annular chamber 31, 31A, 31B, 32 comprises a pair of annular chambers 31, 32. Specifically, the pair of chambers 31, 32 includes a first annular chamber 31 defined by the outer shell 11 and between the trailing edges 101A, 101B of the wings 10A, 10B and the main spar 201. Furthermore, the pair of annular chambers 31, 32 includes a second annular chamber 32 defined by the outer shell 11 and between the main spar 201 and the leading edges 102A, 102B of the wings 10A, 10B. According to another embodiment, two annular chambers 31A, 31B may be provided instead of annular chamber 31. In other words, in this configuration, the inflatable body 10 includes three annular chambers 31A, 31B, and 32. Annular chamber 31A is defined and confined by the outer shell 11 between the trailing edges 101A and 101B of the wings 10A and 10B and the intermediate spar 203, while annular chamber 31B is defined and confined by the outer shell 11 between the intermediate spar 203 and the main spar 203. According to an embodiment, the annular chambers 31, 31A, 31B, and 32 are each inflated to a pressure lower than that at which the at least one main spar 201 and, if provided, the at least one auxiliary spar 202-204 are inflated.

[0046] Reference Figure 2According to an embodiment, both the arched wing 10A and the straight wing 10B include a plurality of planar ribs 206 made of fabric, the ribs 206 intersecting with the at least one main wing spars 201. Specifically, according to an embodiment, the ribs 206 intersect with both the at least one main wing spars 201 and the at least one auxiliary wing spars 202-204. According to an embodiment, the ribs 206 are configured to allow gas to pass through the at least one annular chamber 31, 31A, 31B, 32. In particular, if a plurality of annular chambers 31, 31A, 31B, 32 are provided, for example, two annular chambers 31, 32 or three annular chambers 31A, 31B, 32, then the ribs 206 are configured to allow gas to pass through each of these annular chambers 31, 31A, 31B, 32.

[0047] According to one embodiment, the outer shell 11 includes a radially outer annular portion 11A and a radially inner annular portion 11B. According to one embodiment, ribs 206 connect the inner annular portion 11A and the outer annular portion 11B to allow the arched wing 10A and the straight wing 10B to form a predetermined aerodynamic profile in a closed-wing operating configuration. According to one embodiment, this aerodynamic profile of the wings 10A and 10B particularly has a lens shape. According to one embodiment, the annular portions 11A and 11B are connected to the at least one main spar 201, and more preferably also to the at least one auxiliary spar 202-204.

[0048] refer to Figure 8 According to one embodiment, the aircraft 1 includes a control element 40 associated with at least one annular chamber 31, 31A, 31B, 32 and configured to change the curvature of at least one of the wings 10A, 10B. According to one embodiment, the control element 40 is specifically associated with chamber 31 or chamber 31A. Preferably, the control element 40 is configured to change the curvature of both wings 10A and 10B by modifying the curvature of corresponding control surfaces 80A, 80B. According to one embodiment, the control element 40 includes a soft robotic actuator 40.

[0049] refer to Figure 3 According to an embodiment, the inflatable body includes a plurality of adjacent segments 90, each of the plurality of segments 90 being defined between a pair of consecutive ribs 206 therebetween. More particularly, for example in Figure 3 As can be seen from the embodiments, the arched wing 10A, the straight wing 10B, and the connecting portions 10C and 10D each include multiple adjacent segments 90. According to the embodiments, each segment 90 includes a portion of the outer shell 11, a portion of the at least one main wing spars 201, a portion of the at least one auxiliary wing spars 202-204, and a columnar member 205, which will be described in more detail below. Figure 3In the example, but without introducing any limitations, the inflatable body 10 comprises 36 segments 90. Typically, during the design phase of the aircraft 1, the percentage contribution of each wing 10A, 10B to both aerodynamic and static thrust relative to the total thrust required to balance the weight can be adjusted by changing the distribution of the segments 90 of the domed wing 10A and the straight wing 10B. Different optimal solutions can be obtained depending on the payload 60, flight altitude, and specified cruise speed.

[0050] Reference Figure 2 According to an embodiment, the arched wing 10A and the straight wing 10B each include a plurality of columnar members 205 arranged laterally relative to the trailing edges 101A, 101B and the leading edges 102A, 102B of the arched wing 10A and the straight wing 10B, respectively. Each columnar member 205 is connected to the trailing edges 101A, 101B and the leading edges 102A, 102B of the respective wing 10A, 10B. Furthermore, each columnar member 205 is connected to the at least one main spar 201 and the at least one auxiliary spar 202-204. Advantageously, the columnar members 205 space the at least one main spar 201 and the at least one auxiliary spar 202-204 to provide a structure that reduces deformation of the outer skin 11 upon compression and allows load transfer from the outer skin 11 to the at least one main spar 201 and the at least one auxiliary spar 202-204. Advantageously, according to an embodiment, one of the plurality of columnar elements, columnar element 205, is an inflatable columnar element 205. According to an embodiment, if the at least one main spar 201, the at least one auxiliary spar 202-204, and columnar element 205 are all inflatable elements, then the elements are in fluid communication with each other and are preferably inflated to the same pressure in a closed wing operating configuration.

[0051] According to the implementation, the aircraft 1 does not have airbags to compensate for gas expansion that occurs with changes in altitude. Compared to known technologies based on airship configurations, this solution advantageously enables a significant reduction in the weight and size of the aircraft 1.

[0052] According to the implementation method, unlike fixed-wing stratospheric platforms, aircraft 1 is capable of carrying a payload of 5 kg to 100 kg. In fact, under current technology, such a payload weight level is unacceptable for a fixed-wing configuration due to the insurmountable structural and aeroelastic problems associated with the wide wingspan inherent in fixed-wing stratospheric platforms.

[0053] It should be noted that, according to another embodiment, the aircraft 1 is suitable for use in the troposphere and connected to the ground via a cable of appropriate size. In other words, according to the embodiment, the aircraft 1 is suitable for use as a so-called tethered platform.

[0054] In this regard, it should be noted that currently available tethered airstatic platforms can generate greater aerostatic force than is required to balance the total weight, thereby reducing vertical and horizontal displacement in windy conditions. To generate this additional aerostatic force, the volume of currently available tethered airstatic platforms is larger than that required to balance the total weight.

[0055] The proposed tethered version of the aircraft 1 can counteract the wind by generating aerodynamics. Since no additional ratio of aerostatic force is required, the size of the aircraft 1, and therefore its total weight, is reduced given the same payload and wind conditions.

[0056] The structure of aircraft 1 has been described. Now, considering the case that aircraft 1 is a stratospheric platform, the operating mode of such aircraft will be briefly described by means of non-limiting examples.

[0057] It should be noted that the geometry of aircraft 1 is adapted to takeoff from the initial minimum volume of ground in a predetermined manner ( Figure 9A ) change to hybrid aerodynamic shape (e.g. Figure 9C or Figure 3 Specifically, at takeoff, aircraft 1 has an elliptical shape (…). Figure 9A Furthermore, the weight of the aircraft is balanced solely by static aerodynamic thrust. In this way, takeoff is vertical, requiring no aerodynamic thrust. It should be noted that this vertical takeoff mode allows for takeoff from surfaces that are never prepared; this is a significant advantage for tactical applications. Therefore, by selecting an uncontrolled or at most partially controlled takeoff phase, an initial phase will occur, in which the gas expands until it completely occupies the available volume in at least one chamber 31, 31A, 31B, 32 of the inflatable body 10. The shape of the aircraft 1 evolves from an ellipsoidal shape to a design shape, namely a closed-wing operating configuration. Figure 9C or Figure 3 And continuous transformation. In Figure 9B The diagram shows the configuration of aircraft 1, which is intermediate between the takeoff configuration (…). Figure 9A ) and design and construction ( Figure 9C Between ), the aircraft 1 adopts a closed-wing configuration at a predetermined altitude (ranging from 5000m to 10000m depending on its size) where the aircraft 1 can generate aerodynamics. Figure 9C Once the aerodynamic shape is obtained, i.e., the closed-wing operating configuration (e.g.) Figure 3 or Figure 9CThe second part of the ascent phase begins, at which point static thrust and aerodynamic thrust generated by the aircraft 1 moving forward at a predetermined speed will be utilized. During this phase, the expansion of the gas is not structurally restricted, and excess gas will be released, with lift balancing the portion of the weight that is no longer supported by static airflow.

[0058] Upon reaching the desired flight altitude (e.g., 16,000 m to 21,000 m), aircraft 1 will begin its mission. Two flight modes can be selected depending on the wind strength. The first mode is a ground-stationary mode relative to the area of ​​interest, which is possible when wind speeds are in the range of 7 m / s to 25 m / s. In this mode, the platform utilizes the relative wind to generate lift. For lower wind intensities, aircraft 1 must move along a trajectory, which can be circular or straight, and will allow the necessary lift to be generated to compensate for the static thrust. Obviously, aerodynamic thrust can also compensate for the static thrust loss associated with gas leakage, which is unavoidable in helium-filled systems designed to remain airborne for extended periods. During the descent phase, the internal pressure loss due to altitude will be compensated by introducing air into the outer shell 11 in a controlled and continuous manner to maintain the necessary shape to support and control the platform's descent to the landing surface. Furthermore, during this phase, the system will be assisted in control by electric motors 51 and 52, which utilize energy reserves stored in batteries.

[0059] Aircraft 1, with its closed-wing operating configuration, does not require a wide wingspan. Furthermore, as described above, according to the embodiment, aircraft platform 1 primarily comprises inflatable elements known to be less sensitive to aeroelastic problems, which are associated with composite structures. Unlike conventional airship configurations, the configuration of this invention does not have airbags, thereby significantly reducing the platform's volume and size.

[0060] Based on the foregoing, it can be understood how the aircraft according to this specification can achieve the aforementioned objectives.

[0061] Without prejudice to the principles of the invention, the implementation methods and manufacturing details may vary extensively with respect to the above description disclosed by way of non-limiting examples, without departing from the scope of the invention as defined in the appended claims.

Claims

1. An unmanned hybrid-powered aircraft (1) configured to combine and optimize aerostatic and aerodynamic operations, the aircraft (1) comprising an inflatable body (10) including an outer shell (11) and a load-bearing structure (20) inside the outer shell (11), the inflatable body (10) being adapted to adopt a closed-wing operating configuration; in, In the closed-wing operating configuration, the inflatable body (10) includes: a first arched portion adapted to define an arched wing (10A) having a leading edge (102A) and a trailing edge (101A); a second straight portion adapted to define a straight wing (10B) having a leading edge (102B) and a trailing edge (101B); a third connecting portion (10C) and a fourth connecting portion (10D), the third connecting portion (10C) and the... A fourth connecting portion (10D) is arranged at the opposite end of the straight wing (10B), and the third connecting portion (10C) and the fourth connecting portion (10D) are adapted to join the arched wing (10A) and the straight wing (10B) to each other, wherein the load support structure (20) is an annular structure that extends through the arched wing (10A), the straight wing (10B), and the third connecting portion (10C) and the fourth connecting portion (10D). The load support structure (20) includes at least one annular main wing beam (201). In the closed-wing operation configuration, at least one of the main wing spars (201) includes an arched portion (201A) and a straight portion (201B) of the main wing spar. The arched portion (201A) of the main wing spar is a tubular portion that tapers from the middle of the arched portion (201A) toward the third connecting portion (10C) and the fourth connecting portion (10D). The straight portion (201B) of the main wing spar is a tubular portion that tapers from the middle of the straight portion (201B) toward the third connecting portion (10C) and the fourth connecting portion (10D).

2. The aircraft (1) according to claim 1, wherein, The aircraft (1) is a stratospheric platform (1).

3. The aircraft (1) according to claim 1 or 2, wherein, At least one of the main wing spars (201) is an inflatable wing spars (201).

4. The aircraft (1) according to claim 1, wherein, The load support structure (20) includes at least one annular auxiliary wing beam, the auxiliary wing beam having a cross-sectional dimension smaller than that of the main wing beam (201).

5. The aircraft (1) according to claim 4, wherein, At least one of the auxiliary spars is an inflatable spar.

6. The aircraft (1) according to claim 4, wherein, At least one of the auxiliary spars includes a leading edge spar (204), a trailing edge spar (202), and a middle spar (203). The leading edge spar (204) is disposed at the leading edge (102A) of the arched wing (10A) and the leading edge (102B) of the straight wing (10B). The trailing edge spar (202) is disposed at the trailing edge (101A) of the arched wing (10A) and the trailing edge (101B) of the straight wing (10B). The middle spar (203) is positioned between the main spar (201) and the trailing edge spar (202).

7. The aircraft (1) according to claim 4, wherein, The arched wing (10A) and the straight wing (10B) each include a plurality of columnar members (205) arranged laterally relative to the leading edge (102A) of the arched wing (10A) and the leading edge (102B) of the straight wing (10B), as well as the trailing edge (101A) of the arched wing (10A) and the trailing edge (101B) of the straight wing (10B). Each columnar member (205) is connected to the leading edge (102A) and trailing edge (101A) of the corresponding arched wing (10A) and the leading edge (102B) and trailing edge (101B) of the corresponding straight wing (10B), and each columnar member (205) is connected to at least one main spar (201) and at least one auxiliary spar.

8. The aircraft (1) according to claim 1, wherein, The inflatable body (10) includes at least one annular chamber (31, 31A, 31B, 32) defined by the outer shell (11), and the annular chamber (31, 31A, 31B, 32) extends into the arched wing (10A) and into the straight wing (10B). The aircraft (1) includes a control element (40) associated with at least one of the annular chambers (31, 31A, 31B, 32) and configured to modify the curvature of at least one of the arched wing (10A) and the straight wing (10B).

9. The aircraft (1) according to claim 1, wherein, The leading edge (102A) of the arched wing (10A) and the leading edge (102B) of the straight wing (10B) are offset from each other in a bottom view of the plane of the aircraft (1) in the closed wing operating configuration.

10. The aircraft (1) according to claim 1, the aircraft (1) comprising a propulsion system (51, 52) comprising a pair of mutually spaced forward thrusters (51) associated with the arched wing (10A) and a pair of mutually spaced forward thrusters (52) associated with the straight wing (10B).

11. The aircraft (1) according to claim 1, the aircraft (1) comprising a pair of payload bays (60) located at the third connecting portion (10C) and the fourth connecting portion (10D).

12. The aircraft (1) according to claim 1, wherein, The geometry of the aircraft (1) is adapted to change from an initial ground takeoff shape with minimal volume to a hybrid and aerodynamic shape in a predetermined manner.

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