Field of the invention

[0001] The present invention concerns a stratospheric drone and a method for making and dimensioning a stratospheric drone.

Description of related art

[0002] It has already been suggested to use stratospheric drones as an alternative to satellites for various missions, such as surveillance, weather monitoring, and telecommunication to name just a few. Stratospheric drones are large, unmanned, ultralight, propeller-driven aircraft designed to fly during long periods, usually months or years, just above the tropopause, for example at an altitude between 10 and 30 km. Flying reliably non-stop for months in the lower stratosphere, above severe weather and clouds, requires a new class of aircraft design.

[0003] To achieve long endurance flights, stratospheric drones must use renewable energy. During the day, solar radiation as a source of energy is guaranteed (clouds are rare in the stratosphere). Therefore, most

stratospheric drones are solar-powered, using photovoltaic cells on the wings and/or fuselage. The excess of energy received during the day is stored on board, for example using batteries, and used to power the drone during the night. It has also been suggested to store energy by flying at a higher altitude during daylight hours, and gliding down when the photovoltaic cells stop producing current, thus storing potential energy. This 24-hour day-night cycle offers perpetual flight for aircraft that have been specifically designed for zero-fuel operations.

[0004] Various examples of stratospheric drones are known. For instance, US2014191893 discloses methods and apparatuses for providing wide-area surveillance with a radar and/or other sensors from a stratospheric balloon launched from a land or ship platform for detection, tracking, and

classification of maritime, land, and air objects such as ships, people/vehicles, or aircraft. In one or more embodiments, an apparatus is battery-operated and includes a stratospheric balloon that is filled with helium when it is launched and a gondola with a radar system and communication equipment suspended therefrom. When launched, the apparatus can travel with the wind until it reaches an altitude of approximately 68,500 ft., then it can move substantially horizontally with the stratospheric winds until it returns to earth via a parachute. Multiple apparatus launches at periodic intervals can help provide continuous coverage of the surveillance area. The apparatus can be recovered and re-used or can be considered expendable.

[0005] US2012168555 relates to an autonomous stratospheric aircraft, which is lighter than air, and to a method for providing radio and optical communication, television broadcasting and monitoring with the aid of communication equipment in the aircraft. The disclosed aircraft can be used for producing lighter-than-air aircraft as well as global and regional

communication and television broadcasting and multi-aspect monitoring systems and networks.

[0006] US2013146703 describes an autonomous stratospheric unmanned airship with an operating altitude from 5-22 km and with a multi-month operational cycle. A spheroid rigid geodesic frame of constant volume is formed by a multitude of struts, with an outer envelope enclosing the frame defining the frequency spectrum of the airship above 20 Hz, with vibrational amplitudes between 0.1 and 1 cm. Independently controllable electrical propulsion units, attached to the frame in the horizontal plane passing through the centre of mass, can change the direction and value of the thrust vector. Buoyancy is controlled with a system integrated inside the geodesic frame including buoyant fluid pressurized tanks, valves for the release of the buoyant fluid through the buoyant fluid conduit into the buoyant gas cell, which fills the geodesic frame. Valves at the subsystem support platform enable ambient atmosphere to fill the internal volume of the frame not occupied by the buoyant gas cell.

[0007] Such lighter than air drones have several drawbacks such as exposure to turbulence and high altitude winds during ascent, and low speed.

[0008] JP2009179321 provides an aircraft, configured to have a wide range of flight speeds, consuming low levels of power for an extended period of time, while supporting a communications platform with an unobstructed downward-looking view. The aircraft includes an extendable slat at the leading edge of the wing, and a reflexed trailing edge. The aircraft comprises a flying wing extending laterally between two ends and a centre point. The wing is swept and has a relatively constant chord. The aircraft also includes a power module configured to provide power via a fuel cell. The fuel cell stores liquid hydrogen as fuel, but uses gaseous hydrogen in the fuel cell. A fuel tank heater is used to control the boil-rate of the fuel in the fuel tank. The aircraft includes a support structure including a plurality of supports, where the supports form a tetrahedron that affixes to the wing. Such aircraft is not adapted for perpetual flight.

[0009] US4697761 describes a high altitude solar ultra-lightweight aircraft. The wing span of the rear wing is approximately 120 feet (appr 36,5 m), that of the front wing is approximately 72 feet. The total aircraft weight is 1 160 pounds while the mass of the payload is 100 pounds. This relatively heavy aircraft cannot fly permanently at higher latitudes, since even the energy produced by state-of-the-art photovoltaic cells and by the Long Wave Infrared Radiation cells would not be sufficient to carry such a heavy aircraft during the long winter nights at high latitudes. Moreover, this aircraft comprises one single propeller and one single motor, and is therefore not reliable.

[0010] US2009/026316 describes another solar powered aerial vehicle with a rotatable solar panel. This document does not indicate where the energy storage system could be housed; the only available space seems to be at the nose end of the fuselage, resulting in a very unbalanced construction.

[0011] US201 1/031354 describes another solar powered aircraft. It comprises batteries housed within the plane's main structure. The position of the payload is not indicated.

[0012] WO2015/050609 is related to a method for bonding solar cells to a substrate of a high-altitude aircraft.

[0013] US8448898 describes another unmanned long endurance solar aircraft with a bow shaped wing. [0014] Most of these stratospheric drones have been designed for a specific task, such as transporting a specific telecommunication relay at a given altitude and latitude. These specific requirements result in highly specialized vehicles, well adapted to the specific task, but less suited for transporting different payloads or for navigating at a different altitude.

[0015] There is therefore a need for a more versatile stratospheric drone, in particular a drone adapted for carrying different payloads in a range of different weights and/or for flying at different stratospheric altitudes and/or latitudes.

[0016] While solar powered stratospheric drones are known, a real challenge of a versatile drone is the appropriate overall sizing.

[0017] If we take cars for example, it is apparent that most cars have very similar dimensions; nearly all currently available cars have a width between 1 ,50 and 2,00 m, and a mass between 700 and 3000 kg. This relatively limited range fulfils most needs for most users, so that manufacturers have stopped producing cars outside of this range.

[0018] Surprisingly, there is no consensus yet as to the appropriate size and mass of a solar drone. Various very large stratospheric drones with

considerable payload capabilities have been suggested. Most of these are too costly in development and risky to operate. Also, some of these drones have unrealistic weight targets that are unlikely ever to be achieved.

[0019] While large wing span and high aspect ratio wings have the advantage of increasing aerodynamic efficiency and providing more surface for solar cells, this poses other challenges, for example: aero-elastic

difficulties, increased system complexity, and limited ground operations (runway and airport obstacles). A series of complex trade-offs are required.

[0020] A better understanding of some of the above-mentioned relations is thus required in order to design versatile stratospheric drones.

[0021] In particular, it is an aim of the present invention to propose a method for making a stratospheric drone operating at an altitude between 17 km and 23 km International Standard Atmosphere (ISA). [0022] Another aim is to propose a design which is more versatile and can easily be adapted to various tasks and applications of the drone.

Brief summary of the invention

[0023] According to one aspect, the design of the versatile stratospheric drone is determined by first defining the mass of the payload one wants to transport.

[0024] According to an aspect of the invention, these aims are also achieved by means of a stratospheric drone with a span between 32m and 40m and able to operate at an altitude between 17 km and 23 km ISA, said stratospheric drone comprising:

one fuselage;

at least two wings;

at least one tail;

at least one motor and one propeller;

at least one accumulator in each of said wings, so as to compensate partly the lift force on said wings with the weight of the accumulators in said wings;

photovoltaic cells for charging the accumulators,

a payload pod at the front of the fuselage, for housing a removable payload,

wherein the structure of the drone is adapted for transporting said payload having a mass between 20 and 50 kg.

[0025] This design provides many benefits to increase the versatility of the drone.

[0026] In one aspect, the accumulator in the wings free up space in the fuselage, which becomes available for the payload.

[0027] In another aspect, the weight of the accumulator in the wings compensates at least partly the lift force on the wings. Therefore, the internal bending moments in the wing are reduced, allowing for a lighter

construction. [0028] The two propellers, or more than two propellers, can be mounted to the wings. This frees up space at the nose end of the drone that would be required by a single propeller.

[0029] In a preferred embodiment, the drone further comprises a mechanical and electrical interface between said fuselage and said pod, adapted for the interchangeable connection of different pods to said fuselage. This further increases the versatility of the drone, since different pods with different lengths and/or shape can be provided for housing different type of payloads.

[0030] The different pods that can be connected preferably comprise a means for shifting the centre of gravity of the payload, so as to maintain the position of the aircraft centre of gravity unchanged even when exchanging or replacing payloads.

[0031] In a preferred embodiment, the interchangeable pod is the nose of the drone. This position is advantageous since it is easily accessible. Moreover, it is easier to cool the payload in the pod with air if this payload needs to be cooled down, in particular during the ascent of the drone.

[0032] Moreover, the position of the pod and payload in the front of the fuselage compensates for the weight of the tail.

[0033] The payload in the pod preferably includes electronic components that can be utilized to heat up the volume inside the pod. Therefore, the temperature of the air inside the pod volume can be controlled by in-taking cold external air flow, or heating up with heat emitted from electronic components.

[0034] According to one aspect, the structure of the drone is adapted for transporting a payload having a maximal mass between 20 and 50 kg;

and the total mass (m _to tai) of the stratospheric drone is a multiple of said payload mass, said multiple being in a range between 6 and 15, and determined by the span (b) and the following function: mtotai = a1 * b ⁿ + a2 where a1 = 0.04

a2 = 0.8

2.4 < n < 2.5 for a span b between 32m and 40m.

[0035] The total mass of the drone may be between 190 kg and 440 kg.

[0036] The stratospheric drone is adapted for flight durations, with full control of the drone flight direction and the drone operating systems, depending on latitude, up to 1 month, and preferably up to several months.

[0037] Such a drone, with capabilities to remain fully controlled for a very long duration, preferably several months or years, will be further designated in this document as perpetual flight drone. Such perpetual flights are required for tasks such as remote sensing & photogrammetry,

telecommunication, persistent surveillance, etc.

[0038] Market and technology studies have shown that a payload weight budget between 20 and 50 kg is required to address a variety of stratospheric drone applications and services.

[0039] The multiple k used to determine the total mass of the drone based on the payload weight is determined by experience and statistical data.

[0040] Flight endurance is increased by reducing the power consumption and increasing energy storage. The reduction of power consumption is highly dependent on reducing the overall aircraft weight. However, the design must be economical, robust enough to fulfil the flight mission and to meet the safety requirements of the permit to fly. Within the total weight budget, it is desirable to maximize the payload and accumulator budgets.

[0041] Tests and simulations have shown that the above-indicated relation between mass and span is optimal for such multi-purpose drones, adapted to perform at different altitudes, and under different weight and balance requirements. [0042] The mass of said at least one accumulator is preferably at least 25% of said total mass.

[0043] The invention further provides a method for making a stratospheric drone operating at an altitude between 17 km and 23 km (ISA) for carrying a payload, said stratospheric drone comprising a fuselage, wings, a tail, at least two motors, at least two propellers, a payload pod at the nose of the fuselage, and an accumulator, the method comprising the steps of:

selecting a mass (m _P ayioad) of said payload that the drone needs to carry, said mass being higher than 20 kg and lower than 50 kg;

determining the total mass (m _to tai) of the stratospheric drone as a multiple (k) of said payload mass, said multiple being in a range between 6 and 15, said total mass being determined by the span (b) and the following function:

mtotai = a1 * b ⁿ + a2 * b

where a1 = 0.04, a2 = 0.8, 2.4 < n < 2.5 for a span b between 32m and 40m;

building a drone with said total mass and said span.

[0044] The method advantageously further comprises the steps of:

selecting a ratio between a drag coefficient (CD) and a lift coefficient (CL) of the wings;

determining the power requirement dependent on said drag coefficient, said lift coefficient, the aspect ratio, said span (b), said weight and the air density at said altitude;

determining the accumulator power which is necessary to provide at least this power requirement at night;

determining the mass of the accumulator in order to deliver this accumulator power during a night.

[0045] The method also preferably further comprises the following step:

determining the potential energy recovered by losing altitude at night;

taking this energy into account for determining said accumulator power.

[0046] In a preferred embodiment, the ratio between mass and wing area is between 2 and 5 kg/m ² . [0047] The physical modelling of the stratospheric drone is thus reliant on available statistical data in order to develop new and versatile stratospheric drones with new, previously not considered relations between mass and wingspan.

Brief Description of the Drawings

[0048] The invention will be better understood with the aid of the description of an embodiment given by way of example and illustrated by the figures, in which:

• Figure 1 shows a 3D view of an example of a drone according to the invention;

Figure 2 shows a 3D view of the front part of the drone with two examples of interchangeable payload pods that can be connected to the fuselage.

Figure 3 is a schematic diagram indicating the mass percentage for the main drone components/systems;

Figure 4 shows graphs indicating the relation between altitude, power and energy;

Figure 5 is a schematic diagram showing an example of a system architecture of the aircraft and related efficiencies;

Figure 6 presents a flight diagram which links all kinds of flying objects on one graph;

Figure 7 presents the relationship between mass and span for the drone of the present invention.

Detailed Description of possible embodiments of the Invention

[0049] Figure 1 illustrates an example of an embodiment of a stratospheric drone 1 according to the invention. Such a drone is designed to operate at an altitude between 17 km and 23 km (ISA). [0050] The illustrated drone comprises one fuselage 2, two wings 3 and a tail 4. Those components preferably comprise carbon fibres. The wings are reinforced by a carbon tube forming a wing spar.

[0051] At least two electric motors 5 are provided for driving a

corresponding number of propellers 50. At least one accumulator 60 (Figure 5) is mounted in a span loading manner in the wings and/or in the fuselage. The accumulator may be any device for storing the energy used by the motors 5; it may comprise a battery for storing electricity, or a tank for storing one or a plurality of fuels, such as hydrogens and oxygen, a hydrogen battery, etc. A plurality of accumulators may be interconnected, for example in parallel, in series, or in order to move one fluid from one tank to another one and to balance the weight.

[0052] Photovoltaic cells 7 are provided on the main surfaces of the aircraft. In this example, the photovoltaic cells are provided on the wings top surfaces and both the vertical and horizontal stabilizers. The cells are used to provide power for the electrical motors and all circuits and systems requiring electrical power. The cells' architecture is adapted to provide sufficient power for day operation in addition to a storage capacity enabling night flight.

[0053] The fuselage comprises a pod, such as a hold 6, for housing a removable payload 8 (Figure 2) in a housing 60. The payload pod may also be used to accommodate part of the accumulator system, for example an additional battery and/or a fuel tank.

[0054] The drone could transport different types of payloads with different sizes and different dimensions for different applications. For example, a stratospheric drone could transport a camera, a

telecommunication equipment, a scientific apparatus, etc.

[0055] To achieve this versatility, the payload 8 within a housing 60 in the pod 6 is removable and can be decoupled from the drone (Figure 2). The pod 6 is arranged for accommodating and connecting different types of payloads with different possible weights, dimensions and functions.

[0056] The payload pod 6 itself is removably connected to the fuselage 2. It makes up the nose of the drone, i.e. its frontal extremity. This position is advantageous since the air that flows around the nose in operation can be used to cool down the inner volume of the pod and the payload within this inner volume. Moreover, the weight of the payload compensates for the weight of the tail 4.

[0057] Electrical and mechanical connection means are provided in the pod. In the example, the mechanical connection comprises one flange 20 of the fuselage 2 to which a corresponding flange 65 of the payload pod can be mechanically connected, for example through at least three bolts through holes 61 , 21. At least one electrical connector 22 within the fuselage and a corresponding connector 64 within the pod could be provided for a power and data interface with the pod or with the payload 8 within the pod 6.

[0058] Therefore, various pods 6 with various lengths, shapes and/or inner volume could be provided and connected to a single stratospheric drone for transporting various payloads during different missions.

[0059] The payload preferably comprises electronic components that dissipate heat that could be used for heating the inner volume 60.

[0060] Two examples of different interchangeable pods 6 are illustrated on Figure 2, which differ mainly by their length and inner volume 60 of the payload fairings 62 and 63.

[0061] Some pods may comprise a window for capturing images from a camera inside the pod.

[0062] The pod may also comprise suitable electrical and mechanical connections means for connecting a payload. An accumulator may be placed inside the payload pod to extend the capacity of the wing mounted

accumulator system.

[0063] The different pods 6 comprise means for mounting the payload and/or the accumulator at different positions within the pod, and thus makes sure that the centre of gravity of the drone is not changed when a different pod and different payload is connected. [0064] Figure 4 graphs illustrate the direct relation between altitude (Fig. 4A), power (Fig. 4B) and energy (Fig. 4C) over a period of 80 hours. In Fig. 4B, the line P _PV represents the power delivered by the solar cells; P _acc is the power delivered by the accumulator, P _mec h is the mechanical power transferred to the airflow, and PExcess is the residual power available. Fig. 4C illustrates the energy available in the accumulator (E _acc ) and the residual energy (E _Ex cess)-

[0065] As can be seen, to enable perpetual flight of the drone, night flight must be considered. During that period, energy accumulated during the daytime is used to provide the required power. Thus, flying at night requires energy storage. Two sources of energy are available: potential energy

(altitude) and chemical energy (accumulators). Gliding with no engine power and consuming potential energy does not provide sufficient energy to cover the entire night remaining above the tropopause. Therefore, accumulators remain an important source of energy, as shown in the graphs.

[0066] We will now determine the required power consumption.

[0067] At steady level flight, the lift force generated by the forward motion through the air compensates for the weight (mg), and the propeller thrust (T) compensates for the drag force. In a state of equilibrium, the following equations apply, S being the wing surface and v being the airspeed:

[0068] We can isolate the airspeed v from the first equation: nd then substitute it in the second equation in order to calculate the power r level flight P _mec h: [0069] The speed of the drone is preferably maintained constant, within a narrow range, to optimize the power consumption.

[0070] Using the definition of aspect ratio AR=b2/S where b is the wingspan (or simply span) and S the wing area, we rewrite the previous equation. Therefore, the power required to maintain level flight is:

[0071] This formula describes the power requirement dependent on:

- aerodynamic parameters (drag coefficient, lift coefficient)

- air density at the flight altitude

- wingspan, and aspect ratio

- aircraft total mass.

[0072] Figure 5 is used to represent the internal system architecture of the aircraft and related efficiencies. 5 is the motor(s), 50 are the propellers, 51 the motor controller, 60 the accumulator, 7 corresponds to the solar cells, 70 is the DC/DC and/or DC/AC converter, 8 represents the payload, 81 a DC/DC step- down converter, and 9 the autopilot system. The efficiency of each

component is represented by the symbol η, the power with P.

[0073] The mechanical power P _mech describes the power transferred from the propeller to the environment. This is the available propeller power.

[0074] During the day, the power source is provided by the solar cells which deliver a power sun P _PV . There is an abundance of power most of the time, so the focus is put on the night power provided by the accumulators.

Pynech day ^~ VpropVmotVgearVctrl ^~ ~ iPautopilot+Ppayloacl) [0075] The energy E _acc stored in the accumulators depends on the mass of the accumulators m _aC c multiplied by the coefficient k _acc - During the night power is obtained from the daytime excess energy stored in the accumulators.

'mech day ^~ VpropVmot gearVctrl iPacc

'lbec

3 ^c acc ^li acc" ⁱ acc

acc ^~ γ 'Idischrg ^~ _T 'Idischrg

[0076] In a preferred embodiment, the accumulator mass represents between 25 and 40% of the total mass m _to tai- Figure 3 illustrates an example of the weight percentage of the main drone components or systems.

Considering the size of the drone, the mass of the structure m _st ruct corresponds substantially to half of the total mass. The second component, considered on a weight basis, is the accumulators (m _aC c), requiring between 25-40% of the total mass. The propulsion system, including the motors, (m _pr0 puision) and the payload (m _pa yio _ac i) both represent about 10 % of the total weight. Each percentage may vary by plus or minus 5%; for example, the mass of the structure m _st ruct may be in a range between 45 and 55% of the total mass mtotai-

[0077] In a preferred embodiment, the mass of at least one said

accumulator is at least 25% of said total mass.

[0078] The accumulators are mounted in and distributed along the wing.

[0079] If the payload is lighter and smaller than the nominal value, the remaining volume in the pod may be occupied by an additional accumulator to power the drone. The drone advantageously comprises mechanical and electrical connecting means for housing this additional accumulator in said pod. The function of this additional accumulator is first to increase the autonomy of the drone, and secondly to ensure correct centre of gravity placement.

[0080] Typically, it is very difficult to predict aircraft weight. It is, however, necessary to obtain an estimate in order to select a suitable propulsion system. [0081 ] Figure 6 presents a flight diagram which links all kinds of flying objects on one graph; the abscissa shows the wing loading Weight/Surface while the ordinate corresponds to the weight of the drone.

[0082] The area on the outer left extreme includes light and mostly solar aircraft. On this figure, area 1 1 concentrates most human powered aircrafts. Curve 12 is the Tennekes curve (W/S = 47W ^1/3 , W = 9.6e ⁶ (W/S) ³ . Curve 13 is the Noth sailplane mean model (W/S=2.94W ^{0 37} AR ^{0 45} with AR=20). Curve 14 corresponds to Noth sailplane top 5% model (W/S=0.59W ^{0 35} AR ^{0 84} with

AR=20). Area 15 contains many unmanned solar airplanes, such as Solaris, Sunrise I and II, Sky Sailor, Solar Splinter, Sunrazor, Sun Surfer I and II, and many others. Area 16 contains many manned solar airplanes, such as Solair I and II, Solar Riser, Solar I, Solar Challenger, lcare II and others. Reference number 17 illustrates the Solar Impulse plane.

[0083] The publicly available information about the light, solar aircraft on the outer left of the diagram of Figure 6 is then used to determine an optimal range for the ratio between span b and total mass m _to tai- This ratio is represented in the graph of Figure 7 for a selection of existing unmanned and manned solar aircraft. Curve 22 represents the lightest total mass that can be achieved for each span with carbon fibres structures; the upper limit, curve 21 , is the maximal weight which might be needed in order to build a solid semipermanent vehicle. The large spread of the statistic is attributed to the individual differences between operational objective, flight altitude, endurance, global layout, construction technique, and reliability

requirements.

[0084] Following a detailed analysis, the following relationship for a versatile and robust stratospheric drone is proposed: mtotai = a! * b ⁿ + a ₂ * b where

Ά _Λ = 0.04

a ₂ = 0.8

2.4 < n < 2.5 [0085] In a preferred embodiment n = 2,45. This relationship between the mass nritotai and the span b is represented by the curve 20 on Figure 7.

[0086] Based on this mass to span ratio, numerous simulations have been performed.

[0087] Cruising altitude is preferably set between 17 km and 23 km (ISA) on nominal missions: 17 km being the bottom border for ATC and weather reasons. 23 km being the limit, as air density drops exponentially: 17km has 12% density of sea level, 23 km has 4% of sea-level density.

[0088] The stratospheric drone might be geostationary (station keeping), might autonomously track a route or scan an area.

[0089] In a preferred embodiment in which the payload is substantially 30 kg, the total aircraft weight 300 kg and the wingspan 36m, flight through the night at a nominal altitude of 17km (ISA) throughout the year at a latitude between 0 and 30° North & South, provided the accumulators can store more than 500Wh/kg.

[0090] In operation, the drone preferably stores energy during the day both in its accumulators, for example in batteries in the wings and optionally in the pod, and as potential energy by flying at high altitude during the day. At night, the energy stored in the accumulators is released, and the potential energy is used by allowing the drone to lose altitude.

[0091] In operation, the drone flies in the lower levels of the stratosphere, for example between 16 and 18km ISA during most of the day, and reaches the mid-levels, for example 23 km ISA, shortly before sunset; this altitude is progressively lost during the first part of the night. Therefore, the drone spends more time in lower levels of the stratosphere than in the mid-levels, and uses less power to fly at this lower altitude.