Unmanned hybrid inflatable aircraft

Abstract

A hybrid unmanned aerial vehicle (1) configured for optimized combined use of aerostatic and aerodynamic forces is provided. The aerial vehicle (1) comprises an inflatable body (10) having an outer hull (11) and a load-bearing structure (20) within the outer hull (11). The inflatable body (10) is configured to assume a closed wing operating configuration.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aircraft.

Background Art

[0002] Unmanned aircraft are known. For example, but not necessarily limited to, in recent years, there has been an increasing interest in unmanned aircraft in both the troposphere and the stratosphere, particularly in stratospheric platforms also called HAPS (High Altitude Pseudo Satellite). This interest is motivated by the analysis of applications that can benefit from the advantages obtained from the use of unmanned aircraft, such as land security monitoring, precision agriculture, telecommunications, environmental monitoring, etc.

[0003] Inflatable aircraft-type platforms in both the troposphere and the stratosphere offer four main advantages compared to composite material platforms or metal platforms, namely, reduction of structural weight, reduction of aeroelastic problems due to the nature of the structure, availability of gases lighter than air, and reduction of the transport volume in the contracted state.

[0004] Unlike satellites, HAPS platforms offer the possibility of observing the earth's surface locally, continuously (i.e., without a substantial revisit time), and at a close distance (i.e., from a much lower altitude than satellites). Thereby, even if a payload with low performance and cost efficiency is installed, an image resolution much higher than that provided by satellite remote sensing is provided.

[0005] HAPS are typically platforms capable of operating at altitudes of approximately 18–20 km above the Earth's surface (low stratosphere), and can usually operate continuously for several months using solar energy. The flight altitude of HAPS (18–20 km) is particularly interesting because it far exceeds the altitude range in which civil air traffic is involved. Therefore, the impact of a HAPS platform on air routes is limited to only the ascent phase to mission altitude and the descent phase to the landing base. Furthermore, from a meteorological perspective, a statistical analysis of current wind conditions shows that wind strength is minimal precisely at this altitude range. In the stratosphere, (unlike the troposphere) there is a temperature profile that increases with altitude. This stabilizes the portion of the atmosphere in that area, preventing the formation of updrafts and turbulence.

[0006] In recent years, several proposals have been made regarding HAPS platforms. These proposals can be classified into three main platform types. The first type is represented by a stratospheric balloon, even from a temporal perspective. The second type is represented by a fixed-wing platform, and finally, the third type is represented by an airship-type platform.

[0007] Stratospheric balloon and airship-type platforms rely solely on aerostatic forces (static forces such as buoyancy) to balance their weight (which is lighter than air), while fixed-wing platforms rely solely on aerodynamic forces (dynamic forces such as lift) (which are stronger than air) to balance their weight. Balloon-type platforms have a higher ratio of payload weight to total weight. On the other hand, such balloon-type platforms typically lack propulsion or other directional control systems, and therefore cannot set a predetermined trajectory; their trajectory is determined by wind conditions at various altitudes.

[0008] On the other hand, fixed-wing platforms are based on a configuration with very large wings extending to improve aerodynamic efficiency and are characterized by a very lightweight structure. For the latter reason, the total weight of fixed-wing platforms is limited, so the usable payload weight is reduced to a low weight (5-25 kg). Furthermore, because the structure of these platforms is very light, aeroelastic problems can occur at low altitudes during the ascent and descent phases.

[0009] On the other hand, airship-type platforms are designed for very large payloads (over 200 kg) and, compared to fixed-wing platforms, have a much larger size and a much larger Maximum Takeoff Weight (MTOW) (over 3000-5000 kg). Such a large MTOW is mainly due to the presence of a baronet (air chamber). The baronet forms an inner shell (second shell) that is almost the same size as the outer shell and is necessary to compensate for the expansion of the gas due to changes in altitude (the volume at high altitudes can be up to 20 times that at sea level). Because stratospheric airships are large, ground operations are difficult, especially in bad weather, and dedicated infrastructure is required for the deployment and evacuation of the airship, making tactical use basically impossible. Therefore, from the analysis of the conventional stratospheric platforms described above, two main aspects can be noted. The first aspect concerns the operation of such platforms, and because their operation is heavily dependent on weather conditions, it is thought that they can only be deployed from a limited number of bases around the world. The second aspect concerns the available payload weight, which is limited to 25 kg or less for fixed-wing platforms, while airship-type platforms are only suitable for payloads exceeding 200 kg.

[0010] The object of the present invention is to make available an alternative type of aircraft to the conventional platform described above. This alternative type of aircraft is characterized by a form of aircraft structure (aerodynamic structure) that can efficiently absorb torsional loads and reduce aerodynamic drag without necessarily requiring a large wingspan. [Overview of the Initiative]

[0011] According to one aspect of the present invention, a further objective (alternative to or in addition to the aforementioned objectives) is to make available an aircraft that can reduce the size, volume, cost, and / or operational drawbacks of the aircraft compared to the above-described platform with reference to the prior art.

[0012] According to one aspect of the present invention, a further objective (alternative to or in addition to the aforementioned objectives) is to make available an aircraft that can solve or at least partially overcome the aforementioned drawbacks of the prior art platform.

[0013] According to one aspect of the present invention, a further objective (alternative to or in addition to the aforementioned objectives) is to enable the use of aircraft that can reduce the weight, dimensions, and cost of the aircraft for the same payload compared to the prior art, and that can also be used for tactical operations.

[0014] According to one aspect of the present invention, a further objective (alternative to or in addition to the aforementioned objectives) is to provide an aircraft capable of improving the reliability of the system compared to the prior art systems described above.

[0015] According to one aspect of the present invention, a further objective (alternative to or in addition to the aforementioned objective) is to provide an aircraft capable of carrying a payload in the range of 5 to 100 kg to meet the requirements of a variety of applications.

[0016] The above and other objectives are achieved by an unmanned hybrid inflatable aircraft, as defined in the most common form in appended claim 1 and as defined in some specific embodiments in the dependent claims. [Brief explanation of the drawing]

[0017] The present invention will be better understood from the following detailed description of its embodiments. The description is made by way of example with reference to the following accompanying drawings and is not limited to any aspect.

[0018] [Figure 1] It is a schematic perspective view of a hybrid inflatable aircraft according to a preferred embodiment, seen from the front, with the aircraft shown in a state without a payload. [Figure 2] It shows a schematic perspective view of the load-bearing structure inside the aircraft of FIG. 1, seen from the front. [Figure 3] It is a front view of the aircraft of FIG. 1, seen from the front, and the payload is also shown. [Figure 4] It is a rear view of the aircraft of FIG. 1, seen from the rear, with the aircraft shown in a state without a payload. [Figure 5] It is a side view of the aircraft of FIG. 1, seen from the side, with the aircraft shown in a state without a payload. [Figure 6] It is a plan view of the aircraft of FIG. 1, seen from above, with the aircraft shown in a state without a payload. [Figure 7] It is a bottom view of the aircraft of FIG. 1, seen from below, with the aircraft shown in a state without a payload. [Figure 8] It is a perspective view showing a part of the aircraft of FIG. 1, with a part of the outer shell of the aircraft broken and shown. [Figure 9A] It is a schematic front view of the aircraft of FIG. 1 in the initial takeoff configuration, seen from the front. [Figure 9B] It is a schematic front view of the aircraft of FIG. 1 in the intermediate configuration, seen from the front. [Figure 9C] It is a schematic front view of the aircraft of FIG. 1 in the operating mission configuration, seen from the front.

DETAILED DESCRIPTION OF THE INVENTION

[0019] The same or equivalent elements in the accompanying drawings are given the same reference numerals.

[0020] Furthermore, in the following description, the terms “down,” “up,” “horizontal,” and “vertical” used to describe parts of the aircraft according to the present invention are intended to refer to the direction of the aircraft in its normal operating state at the assigned mission altitude. In addition, in the following description, the terms “forward” and “rear” used to describe parts of the aircraft according to the present invention are intended to refer to the direction along the aircraft’s direction of travel X1 (see Figure 6) in its normal operating state at the assigned mission altitude. Furthermore, the terms “radially outward” and “radially inward” used to describe parts of the aircraft platform according to the present invention are intended to refer to the radial direction around the axis of the opening 12, which will be described later (see, for example, Figure 3). Here, the axis passes through the center of the opening 12 and is perpendicular to the opening 12.

[0021] First, referring to FIGS. 1 to 7, an unmanned hybrid aircraft according to a preferred embodiment is generally indicated by reference numeral 1. According to one embodiment, in particular, the aircraft 1 is a stratospheric platform, more preferably a HAPS (High Altitude Pseudo Satellite) stratospheric platform. Further, it should be noted that for the purpose of this description, the term "hybrid" used to define the aircraft indicates that the aircraft is configured to utilize an optimized combination of aerostatic forces (static forces such as buoyancy) and aerodynamic forces (dynamic forces such as lift) in a balanced manner. Here, the term "optimized" means selecting a buoyancy ratio (i.e., the ratio of aerostatic force to the total force) to minimize the total weight and installation area of the platform (aircraft) while fixing the payload and minimum altitude achievable by only aerostatic forces (static forces). In particular, the hybrid aircraft according to the present disclosure is configured to utilize both aerostatic (static) buoyancy and aerodynamic (dynamic) lift. In other words, the aircraft 1 is conveniently configured to utilize aerodynamic (dynamic) forces not only for controlling the aircraft 1 but also for balancing the weight of the aircraft itself at various flight stages. According to one embodiment, the aircraft 1 is a tactical aircraft. That is, the aircraft 1 can be easily transported within a standard container and directly deployed into an operational scenario. This is made possible by the small size and MTOW (Maximum Takeoff Weight) of the aircraft 1. Referring to FIGS. 5 and 6 to 7, according to one embodiment, in particular, the aircraft 1 has a maximum length L1 (wingspan L1) in the range of 8 m to 25 m, an MTOW (Maximum Takeoff Weight) in the range of 30 kg to 400 kg, and a payload 60 in the range of 5 kg to 100 kg. Generally, the aircraft 1 is configured to be used for terrestrial surveillance. In particular, the aircraft 1 can be used for a plurality of different applications, although not limited thereto, such as border surveillance, environmental surveillance, precision agriculture, communication, homeland security, emergency support, and the like.

[0022] According to one embodiment, as will be understood in more detail in the following description, the aircraft 1 has a combination of several inflatable structural elements (gas-filled inflatable elements) having different internal pressures. These inflatable structural elements are preferably bonded to rigid structures (made of, for example, composite materials, aluminum, etc.), such as nacelles, engine mounts, and movable control surfaces. According to one embodiment, all inflatable structural elements of the aircraft 1 described below in this disclosure comprise a laminated material consisting of a gas-retaining layer, a structural layer, and a protective layer.

[0023] Referring to Figures 1 to 7, the aircraft 1 comprises an inflatable body (gas-filled body) 10. The inflatable body 10 has an outer shell 11 (skin 11) and a load-bearing structure 20 inside the outer shell 11. Conveniently, the inflatable body 10 is configured to operate in a closed-wing configuration (operational mission configuration) (see Figures 1, 3 to 6, 7, and 9C). In particular, in the closed-wing configuration, the inflatable body 10 has an annular structure extending around a through-opening 12. According to one embodiment, the closed-wing configuration corresponds to a design configuration (operational mission configuration) in which the aircraft 1 is inflated with gas (preferably helium). According to one embodiment, in the closed-wing operating configuration, thanks to its unique aerodynamic structure, aircraft 1 can very efficiently generate both dynamic lift equivalent to approximately 60% to 80% of the total weight of aircraft 1 and aerostatic (static) buoyancy equivalent to approximately 40% to 20% of the total weight of aircraft 1. Here, aircraft 1 is filled with a gas lighter than air, and the gas is preferably helium as described above.

[0024] According to one embodiment, in the closed airfoil configuration, the inflatable body 10 includes an arched first portion 10A configured to define an arched wing 10A having a leading edge 102A and a trailing edge 101A. Preferably, the arched wing 10A is semi-elliptical or substantially semi-elliptical. Furthermore, in the closed airfoil configuration, the inflatable body 10 includes a linear second portion 10B configured to define a linear wing 10B having a leading edge 102B and a trailing edge 101B. In particular, the arched wing 10A is the upper wing, and the linear wing 10B is the lower wing. In the closed airfoil configuration, the inflatable body 10 further includes connecting portions 10C, 10D (a third connecting portion 10C and a fourth connecting portion 10D) located at both ends of the linear wing 10B. The connecting portions 10C, 10D are configured to connect the arched wing 10A and the linear wing 10B to each other. In other words, the connecting parts 10C and 10D correspond to both ends, or tips, of the respective wings 10A and 10B. In fact, thanks to the connecting parts 10C and 10D, the straight wing 10B closes the arched wing 10A both structurally and aerodynamically.

[0025] Referring to Figures 5 and 7, in one embodiment, in a bottom view of the aircraft 1 in the closed-air configuration, the leading edges 102A and 102B are aligned with each other, while the trailing edges 101A and 101B are offset from each other. However, in an alternative embodiment, in a bottom view of the aircraft 1 in the closed-air configuration, the leading edges 102A and 102B may be offset from each other to improve the stability and controllability characteristics of the aircraft 1. In fact, by shifting the pressure centers of the arched wing 10A and the straight wing 10B, for example, the longitudinal moment with respect to the center of gravity can be canceled out.

[0026] According to one embodiment, control surfaces 80A and 80B are associated with and provided on the arched wing 10A and the straight wing 10B, respectively. The control surfaces 80A and 80B perform control and trim functions for the aircraft 1. According to one embodiment, the control surfaces 80A and 80B may be rigid structures made of composite materials.

[0027] According to one embodiment, the aircraft 1 comprises solar panels 70 which may be associated with arched wings 10A and / or straight wings 10B. The solar panels 70 are preferably flexible solar panels 70. According to one embodiment, the aircraft 1 comprises a pair of housings (i.e., nacelles) for a payload 60 (preferably batteries and avionics systems). The housings are located in a third connecting portion 10C and a fourth connecting portion 10D.

[0028] According to one embodiment, the aircraft 1 comprises at least one propulsion system 51, 52. The propulsion system 51, 52 preferably includes a pair of front propellers 51. The pair of front propellers 51 are spaced apart from each other and are provided in association with an arched wing 10A. Furthermore, the propulsion system 51, 52 includes another pair of front propellers 52. The pair of front propellers 52 are spaced apart from each other and are provided in association with a straight wing 10B. However, according to one embodiment, five or more propellers 51, 52 may be provided. According to one embodiment, the pair of front propellers 51 are aligned with each other along the extending direction of the arched wing 10A, and the other pair of front propellers 52 are aligned along the extending direction of the straight wing 10B. For example, as shown in Figures 6 and 7, according to one embodiment, a pair of front propellers 51 are positioned relatively close to each other, while another pair of front propellers 52 are positioned relatively far apart from each other. According to one embodiment, the propellers 51 and 52 include helical propellers. More specifically, according to one embodiment, the propellers 51 and 52 comprise driven helical propellers. Preferably, the propellers 51 and 52 are equipped with electric motors. In particular, according to one embodiment, the propulsion systems 51 and 52 are entirely electric, and power generation is based on the aforementioned solar panels 70. More specifically, according to one embodiment, the aircraft 1 is configured to be completely energy self-sufficient and able to fly for several weeks. Preferably, energy storage necessary to ensure nighttime flight is based on high-energy-density batteries such as lithium polymer batteries, lithium-ion batteries, or lithium-sulfur batteries.

[0029] The specific closed airfoil configuration of aircraft 1 offers significant structural and aerodynamic advantages. Structurally, the arched wing 10A has only a small amount of load distributed on it (partially supported by aerostatic thrust from an internal gas). For example, according to one embodiment, the load distributed on the arched wing 10A is due to the weight of the wing 10A itself, preferably due to the presence of the solar panel 70, the propeller 51, and / or the control element 40 (which will be detailed later with reference to Figure 8). However, the majority of the load is concentrated at the ends (tips) of the wing 10A, i.e., the connection points 10C, 10D where the payload 60 (preferably batteries and avionics systems) is located. According to one embodiment, due to the arched structure and the specific load configuration at the connection points 10C, 10D, the wing 10A must primarily withstand tensile loads. The tensile loads are best suited to the material of the wing 10A itself (preferably mainly fabric). According to one embodiment, structurally, the straight wing 10B will have only the following distributed loads, namely the weight of the wing 10B itself, and preferably the weight of the propeller 52 and / or the control element 40 (described later). The straight wing 10B will contribute only to a portion of the load concentrated at the ends (tips) of the wing 10B, i.e., the connection points 10C, 10D. According to one embodiment, the straight wing 10B mainly needs to support bending loads. These bending loads are mitigated by both aerostatic (static) buoyancy and aerodynamic (dynamic) loads generated by the arched wing 10A at both ends, i.e., the connection points 10C, 10D. From a structural standpoint, the closed airfoil configuration efficiently solves the need to absorb torsional loads typical of classic fixed wings with free ends. Furthermore, such loads are optimally absorbed by the load-bearing structure 20 and the outer shell (skin) 11.

[0030] From an aerodynamic standpoint, the closed airfoil configuration offers an optimal solution for reducing induced drag without necessarily considering a wide wingspan. The aerodynamic (dynamic) load is greater in the arched wing 10A than in the straight wing 10B. Furthermore, due to its arched configuration, the arched wing 10A generates a lateral aerodynamic (dynamic) force in addition to lift, and this force helps to pull the straight wing 10B and support the bending load of the straight wing 10B.

[0031] Referring to Figure 2, according to one embodiment, the load-bearing structure 20 is an annular structure extending through an arched wing 10A, a straight wing 10B, and the aforementioned connecting parts 10C, 10D (third connecting part 10C and fourth connecting part 10D). According to one embodiment, the load-bearing structure 20 comprises at least one annular main spar 201. Advantageously, according to one embodiment, at least one main spar 201 is a spar having an inflatable structure (gas-filled). In this embodiment, the load-bearing structure 20 comprises a single main spar 201. According to one embodiment, in the closed airfoil operating configuration, at least one main spar 201 comprises an arched main spar section 201A and a straight main spar section 201B. The arched main spar section 201A is associated with the arched wing 10A, while the straight main spar section 201B is associated with the straight wing 10B. In particular, the arched main girder section 201A is preferably a tubular section having a circular cross-section, tapering from the center of the arched main girder section 201A toward the connecting third section 10C and the connecting fourth section 10D. Furthermore, the straight main girder section 201B is preferably a tubular section having a circular cross-section. The straight main girder section 201B tapering from the center of the straight main girder section 201B toward the connecting third section 10C and the connecting fourth section 10D. Preferably, the cross-section of the straight main girder section 201B changes in particular as the ratio of the thickness of the contour of the straight wing 10B changes.

[0032] Referring again to Figure 2, according to one embodiment, the load-bearing structure 20 comprises at least one annular sub-girders 202-204 having a cross-section smaller than that of the main girder 201. According to one embodiment, it is advantageous that at least one sub-girder 202-204 is an inflatable (gas-injected) girder. According to one embodiment, at least one sub-girder 202-204 comprises a trailing edge girder 202, a leading edge girder 204, and an intermediate girder 203. In particular, the trailing edge girder 202 is located on the trailing edges 101A, 101B of the arched wing 10A and the straight wing 10B. The leading edge girder 204 is located on the leading edges 102A, 102B of the arched wing 10A and the straight wing 10B. The intermediate girder 203 is provided between at least one main girder 201 and the trailing edge girder 202. According to one embodiment, the intermediate spar 203 may be positioned at the location of the maximum thickness ratio of the wings 10A, 10B. The distance from the trailing edge spar 202 to the main spar 201 is equal to 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 sub-spars 202-204 are inflatable (gas-inflated) spars. According to one embodiment, at least one main spar 201 and at least one sub-spar 202-204 are fluidly connected to each other and preferably inflated to the same pressure.

[0033] Referring again to Figure 2, according to one embodiment, the inflatable body 10 is defined by an outer shell (skin) 11 and comprises at least one annular chamber 31, 31A, 31B, 32 extending into an arched wing 10A and a straight wing 10B. According to one embodiment, the inflatable body 10 comprises, in particular, a plurality of annular chambers 31, 31A, 31B, 32. According to one embodiment, the plurality of annular chambers 31, 31A, 31B, 32 comprises chambers that are in fluid communication with each other. For example, according to one embodiment, the outer shell (skin) 11 is fixed, preferably bonded, only 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, for example, if there are ailerons (auxiliary wings) with chambers, the outer shell 11 may be fixed, preferably bonded, to the intermediate spar 203. According to one embodiment, in the operating mode of a closed airfoil, at least one annular chamber 31, 31A, 31B, 32 is inflated to a pressure lower than the pressure at which at least one main spar 201 and at least one secondary spar 202-204 are inflated. According to a convenient embodiment, at least one annular chamber 31, 31A, 31B, 32 comprises a pair of annular chambers 31, 32. In particular, the pair of chambers 31, 32 includes a first annular chamber 31. The first annular chamber 31 is defined by the outer shell 11 and is defined 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. The second annular chamber 32 is defined by the outer shell 11 and is defined between the main spar 201 and the leading edges 102A, 102B of the wings 10A, 10B. In a further embodiment, instead of the annular chamber 31, two annular chambers 31A, 31B may be provided. In other words, in such a case, the inflatable body 10 comprises three annular chambers 31A, 31B, 32. In this case, the annular chamber 31A is defined by the outer shell 11 and is defined between the trailing edges 101A, 101B of the wings 10A, 10B and the intermediate spar 203, and the annular chamber 31B is defined by the outer shell 11 and is defined between the intermediate spar 203 and the main spar 201.According to one embodiment, the annular chambers 31, 31A, 31B, and 32 are each inflated to a pressure lower than the pressure at which at least one main girder 201 and at least one sub-girders 202 to 204 are inflated.

[0034] Referring to Figure 2, according to one embodiment, the arched wing 10A and the straight wing 10B each comprise a plurality of planar ribs 206 made of fabric. At least one main spar 201 intersects the plurality of ribs 206. In particular, according to one embodiment, both at least one main spar 201 and at least one secondary spar 202-204 intersect the rib 206. According to one embodiment, the rib 206 is configured to allow gas to pass through at least one annular chamber 31, 31A, 31B, 32. In particular, when a plurality of annular chambers 31, 31A, 31B, 32 (for example, two annular chambers 31, 32, or three annular chambers 31A, 31B, 32) are provided, the rib 206 is configured to allow gas to pass through each of such annular chambers 31, 31A, 31B, 32.

[0035] According to one embodiment, the outer shell 11 comprises a radially outer annular portion 11A and a radially inner annular portion 11B. According to one embodiment, the rib 206 connects the outer annular portion 11A and the inner annular portion 11B. This allows for the formation of predetermined aerodynamic contours of the arched wing 10A and the straight wing 10B in the operating mode of a closed airfoil. According to one embodiment, such aerodynamic contours of the wings 10A and 10B have a particularly lenticular shape. According to one embodiment, the outer annular portion 11A and the inner annular portion 11B are connected to at least one main spar 201, and more preferably to at least one secondary spar 202-204.

[0036] Referring 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. The control element 40 is configured to change the curvature of at least one wing 10A, 10B. According to one embodiment, the control element 40 is associated particularly with chamber 31 or chamber 31A. Preferably, the control element 40 is configured to change the curvature of both wings 10A, 10B, preferably by changing the curvature of the corresponding control surfaces (control surfaces 80A, 80B). According to one embodiment, the control element 40 includes a soft robotic actuator.

[0037] Referring to Figure 3, according to one embodiment, the inflatable body 10 comprises a plurality of mutually adjacent segments (sections) 90. Ribs 206 are arranged between adjacent segments 90. Each segment 90 is defined between a pair of adjacent ribs 206. More specifically, as seen in Figure 3, for example, according to one embodiment, the arched wing 10A, the straight wing 10B, and the connecting sections 10C, 10D comprise a plurality of mutually adjacent segments 90. According to one embodiment, each segment 90 comprises a part of the outer shell 11, a part of at least one main spar 201, a part of at least one secondary spar 202-204, and a strut 205, which will be described in more detail later in this specification. In the example of Figure 3, the inflatable body 10 comprises 36 segments 90, although this is not limited to this example. Generally, during the design phase of aircraft 1, by changing the distribution of segments 90 in the arched wing 10A and the straight wing 10B, it becomes possible to adjust the proportion of how much each wing 10A and 10B contributes to both aerodynamic (dynamic) thrust and aerostatic (static) thrust in relation to the total thrust required to balance the weight. Depending on the payload 60, flight altitude, and assigned cruising speed, various optimal solutions can be achieved.

[0038] Referring to Figure 2, according to one embodiment, the arched wing 10A and the straight wing 10B each include a plurality of struts 205. The struts 205 are arranged along the transverse direction with respect to the trailing edges 101A, 101B and leading edges 102A, 102B of the arched wing 10A and the straight wing 10B. Each strut 205 is connected to the trailing edges 101A, 101B and leading edges 102A, 102B of the respective wings 10A, 10B. Furthermore, each strut 205 is connected to at least one main girder 201 and at least one secondary girder 202-204. Advantageously, the struts 205 form a gap between the at least one main girder 201 and at least one secondary girder 202-204. This provides a structure that reduces the deformation of the outer shell 11 by being compressed, and enables load transmission from the outer shell 11 to at least one main spar 201 and at least one secondary spar 202-204. Advantageously, according to one embodiment, the plurality of struts 205 are inflatable struts having an inflatable structure. According to one embodiment, when at least one main spar 201, at least one secondary spar 202-204, and struts 205 are all inflatable elements, it is preferable that these elements are in fluid communication with each other and inflated to the same pressure in the closed airfoil operating configuration.

[0039] According to one embodiment, the aircraft 1 does not have a baronet to compensate for gas expansion due to changes in altitude. The solution of this embodiment advantageously allows for a significant reduction in the weight and size of the aircraft 1 compared to known technologies based on airship-type configurations.

[0040] According to one embodiment, aircraft 1 is capable of carrying a payload of 5-100 kg, unlike fixed-wing stratospheric platforms. In fact, such a payload weight class is not permissible in fixed-wing configurations because conventional technology inevitably presents insurmountable structural and aeroelastic problems associated with wide wingspans in fixed-wing stratospheric platforms.

[0041] In a further embodiment, aircraft 1 is configured to be used in the troposphere and connected to the ground by a cable of appropriate size. In other words, in one embodiment, aircraft 1 is configured to be used as a so-called tether platform (tether satellite system).

[0042] In this regard, currently available tethered aerostat platforms are capable of generating aerostatic forces (static forces) greater than those required to balance the overall weight to reduce both vertical and planar displacements in the presence of wind. To generate such excessive aerostatic forces, the volume of currently available tethered aerostat platforms is larger than the volume required to balance the overall weight.

[0043] The tether configuration of aircraft 1 in this embodiment can counteract wind by generating aerodynamic forces (dynamic forces). Therefore, since no excessive aerostatic forces (static forces) are required, the size of aircraft 1, and consequently its overall weight, is reduced, assuming the same payload and the same wind conditions.

[0044] The structure of aircraft 1 has been described above. Next, considering the case where aircraft 1 is a stratospheric platform, the operating modes of the aircraft described above will be briefly explained as non-limiting examples.

[0045] The geometric shape of aircraft 1 is configured to change in a predetermined manner from an initial minimum-volume takeoff configuration (Figure 9A) to a hybrid aerodynamic design configuration (operational mission configuration) (e.g., Figure 9C or Figure 3). In particular, during takeoff, aircraft 1 has an ellipsoidal shape (Figure 9A), and the weight of aircraft 1 is balanced solely by aerostatic thrust (static thrust). In this way, takeoff is performed vertically without the need for aerodynamic thrust (dynamic thrust). This vertical takeoff mode allows for departure from surfaces that are not specifically prepared. This is a significant advantage for tactical applications. Therefore, in the takeoff phase, which is either uncontrolled by selection or at best partially controlled, there is an initial stage in which gas expands until it completely occupies the available volume in at least one chamber 31, 31A, 31B, 32 of the inflatable body 10. As aircraft 1 ascends, its shape continuously deforms from an ellipsoid shape to its design shape (design configuration), i.e., a closed-wing operating configuration (Figure 9C or Figure 3). Figure 9B shows the intermediate configuration of aircraft 1 between the takeoff configuration (Figure 9A) and the design configuration (Figure 9C). Aircraft 1 adopts the closed-wing operating configuration (Figure 9C) at a predetermined altitude (5000-10000m depending on size) in which aircraft 1 can generate aerodynamic thrust (dynamic thrust). Once the aerodynamic shape, i.e., the closed-wing operating configuration (e.g., Figure 3 or Figure 9C), is acquired, the second part of the ascent phase begins. At this point, both aerostatic thrust (static thrust) and aerodynamic thrust (dynamic thrust) generated by aircraft 1 moving forward at a predetermined speed are utilized. At this stage, the expansion of the gas is not structurally suppressed, and the excess gas is released. The balance of the heavy parts, which would otherwise be unsupported aerostatically, is maintained by dynamic lift.

[0046] Once the required flight altitude (e.g., 16,000–21,000 meters) is reached, aircraft 1 commences its mission. Two flight modes are possible depending on the wind strength. The first mode is stationary relative to the target zone and is possible when the wind speed is in the range of 7–25 m / s. In the first mode, the platform utilizes the relative wind to generate lift. When the wind strength is weak, aircraft 1 needs to move along a trajectory, for example, circular or linear, thereby generating the lift necessary to compensate for aerostatic thrust (static thrust). Naturally, aerodynamic thrust (dynamic thrust) can also compensate for the loss of aerostatic thrust (static thrust) associated with gas leaks, which are unavoidable in helium-filled systems intended to remain airborne for extended periods. During the descent phase, the loss of internal pressure due to altitude is compensated by introducing air into the outer shell 11 in a controlled and continuous manner. This maintains the necessary shape that helps support and control the platform as it descends to the landing surface. Furthermore, at this stage, the control of the system can be supported by electric motors (propulsion systems 51, 52) that utilize energy stored in the battery.

[0047] Aircraft 1 does not require a wide wingspan thanks to its closed-aircraft design. Furthermore, as described above, according to one embodiment, aircraft 1 has inflatable elements as its main components, which are known to be less susceptible to aeroelastic problems associated with composite structures. Unlike conventional airship-type configurations, the present invention does not have a baronet, thus significantly reducing both the volume and size of the platform.

[0048] Therefore, based on the above explanation, it can be understood how the aircraft according to this embodiment can achieve the above-mentioned objectives.

[0049] Without impairing the principles of the present invention, embodiments and details of the manufacture can be broadly modified from the above description disclosed as non-limiting examples, without departing from the scope of the invention as defined in the appended claims.

Claims

1. A hybrid unmanned aircraft (1) configured to optimize and utilize a combination of aerostatic and aerodynamic forces, The inflatable body (10) comprises an outer shell (11) and a load-bearing structure (20) inside the outer shell (11), The inflatable body (10) is configured to operate in a closed-wing configuration. In the closed airfoil configuration, the inflatable body (10) comprises: an arched first portion (10A) configured to define an arched wing (10A) having a leading edge (102A) and a trailing edge (101A); a straight second portion (10B) configured to define a straight wing (10B) having a leading edge (102B) and a trailing edge (101B); and connecting third portion (10C) and fourth portion (10D) positioned at both ends of the straight wing (10B) and configured to connect the arched wing (10A) and the straight wing (10B) to each other. The load-bearing structure (20) is an annular structure that extends through the arch-shaped wing (10A), the straight wing (10B), the connecting third portion (10C), and the connecting fourth portion (10D). The load-bearing structure (20) comprises at least one annular main girder (201), In the closed airfoil configuration described above, the main spar (201) comprises an arched main spar section (201A) and a straight main spar section (201B), The arch-shaped main girder section (201A) is a tubular section that tapers from the center of the arch-shaped main girder section (201A) toward the connecting third section (10C) and the connecting fourth section (10D), The straight main girder section (201B) is a tubular section that tapers from the center of the straight main girder section (201B) toward the connecting third section (10C) and the connecting fourth section (10D). aircraft.

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

3. The main girder (201) has an inflatable structure. The aircraft according to claim 1 or claim 2.

4. The load-bearing structure (20) comprises at least one annular subgirders (202-204) having a cross-section smaller than that of the main girder (201). The aircraft according to claim 1.

5. The aforementioned subgirders (202-204) have an inflatable structure. The aircraft according to claim 4.

6. The aforementioned sub-girders (202-204) are, A leading edge spar (204) is positioned at the leading edge (102A) of the arched wing (10A) and the leading edge (102B) of the straight wing (10B), A trailing edge spar (202) is positioned at the trailing edge (101A) of the arched wing (10A) and the trailing edge (101B) of the straight wing (10B), Including an intermediate girder (203) located between the main girder (201) and the trailing edge girder (202), The aircraft according to claim 4.

7. The arched wing (10A) comprises a plurality of support columns (205) arranged along the transverse direction with respect to the leading edge (102A) and trailing edge (101A) of the arched wing (10A), and connected to the leading edge (102A) and trailing edge (101A) of the arched wing (10A), the main spar (201), and the secondary spar (202-204). The linear wing (10B) comprises a plurality of struts (205) arranged along the transverse direction of the leading edge (102B) and trailing edge (101B) of the linear wing (10B) and connected to the leading edge (102B) and trailing edge (101B) of the linear wing (10B), the main spar (201), and the secondary spars (202-204). The aircraft according to claim 4.

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

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

10. The propulsion system (51, 52) includes a pair of forward propellers (51) associated with the arched wing (10A) and spaced apart from each other, and a pair of forward propellers (52) associated with the straight wing (10B) and spaced apart from each other. The aircraft according to claim 1.

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

12. The geometric shape of the aircraft (1) is configured to change in a predetermined manner from an initial minimum-volume takeoff configuration to a hybrid aerodynamic configuration. The aircraft according to claim 1.

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