CN112498741B - Detection aircraft and Mars cruise detection method - Google Patents
Detection aircraft and Mars cruise detection method Download PDFInfo
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- CN112498741B CN112498741B CN202011193409.1A CN202011193409A CN112498741B CN 112498741 B CN112498741 B CN 112498741B CN 202011193409 A CN202011193409 A CN 202011193409A CN 112498741 B CN112498741 B CN 112498741B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
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- B—PERFORMING OPERATIONS; TRANSPORTING
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Abstract
The invention discloses a detection aircraft, comprising: the wing is in a wing shape and provided with an energy supply system, and detection equipment is arranged on the wing; the magnus rotor is rotatably arranged at the front end of the wing and is arranged parallel to the front end of the wing, and the energy supply system is in transmission connection with the magnus rotor to drive the magnus rotor to rotate; the parafoil is arranged on the upper surface of the wing, and the energy supply system is electrically connected with the parafoil to supply energy to the parafoil. The magnus rotor is installed additional at the wing leading edge through what this application detected aircraft, cruises the in-process, and the magnus rotor can provide bigger lift for the wing at the rotation in-process, and the magnus rotor can provide necessary lift demand for mars detector through measures such as changing the rotational speed under the atmospheric low-density low dynamic pressure condition of outer planet.
Description
Technical Field
The invention relates to the technical field of space exploration, in particular to an exploration aircraft.
Background
At present, three detection methods of the planet mainly comprise surrounding detection, in-situ detection and inspection tour detection. The surrounding detection is mainly completed by a spacecraft running on an orbit surrounding other planets such as mars, Venus and the like by means of remote sensing detection, and if the distance between a detector for surrounding mars detection and the surface of the mars is generally more than hundreds of kilometers, the details on the surface of the mars are not easy to observe clearly. Meanwhile, after a certain place is observed by the satellite once, the place is observed again only after a long time interval due to the limitation of the operation orbit, and the phenomenon with rapid change is difficult to be observed continuously. In-situ detection is to detect the rock, atmosphere or biological information of the detector by using the detector landed on the surface of the planet. The tour inspection is also performed on the surface of a planet, but the detection range can be expanded compared with the in-situ inspection because the tour inspection is performed by a movable tour device such as a mars vehicle. For the Mars detection, although the fineness of the in-situ detection and the patrol detection is higher, the in-situ detection can only be expanded at a landing point, and the walking range of a patrol device such as a Mars vehicle is quite limited. The "opportunity number" mars, which just finished working shortly before, for example, moved only 45 km in total during 14 years of work, which is negligible with respect to the week length of many twenty thousand kilometers of mars.
Disclosure of Invention
The present invention is directed to solving at least one of the problems of the prior art. Therefore, the invention provides a detection aircraft which can carry out cruise detection on planets for a long time.
The invention further provides a mars cruising detection method.
The technical scheme adopted by the embodiment of the invention is as follows: a probe vehicle comprising: the wing is in a wing shape and provided with an energy supply system, and detection equipment is arranged on the wing; the magnus rotor is rotatably arranged at the front end of the wing and is arranged parallel to the front end of the wing, and the energy supply system is in transmission connection with the magnus rotor to drive the magnus rotor to rotate; the parafoil is arranged on the upper surface of the wing, and the energy supply system is electrically connected with the parafoil to supply energy to the parafoil.
The detection aircraft provided by the embodiment of the invention has at least the following beneficial effects: the magnus rotor is installed additional at the wing leading edge through what this application detected aircraft, cruises the in-process, and the magnus rotor can provide bigger lift for the wing at the rotation in-process, and the magnus rotor can provide necessary lift demand for mars detector through measures such as changing the rotational speed under the atmospheric low-density low dynamic pressure condition of outer planet.
According to some embodiments of the invention, the energy supply system comprises a solar cell array and a rechargeable lithium-carbon dioxide cell, the solar cell array covers the outer surface of the wing, the lithium-carbon dioxide cell is arranged in the wing, and the solar cell array is connected with the lithium-carbon dioxide cell to charge the lithium-carbon dioxide cell.
According to some embodiments of the invention, the detection device is arranged inside the wing and is arranged at the middle front position of the wing, the lower surface of the wing adopts a transparent structure, and the detection device detects through the lower surface of the wing.
According to some embodiments of the invention, the wing is provided with a landing gear at its bottom.
According to some embodiments of the invention, the magnus rotor is rotated in a clockwise direction.
According to some embodiments of the invention, the magnus rotor has a ratio of circumferential speed to incoming flow speed of 0-4.
According to some embodiments of the invention, the ratio of the length of the parachute line of the parafoil to the length of the chord of the wing is 1-2.
The invention also discloses a Mars cruising detection method, which comprises the following steps:
step 1: in the ejection starting stage, the detection aircraft is ejected and lifted off through an ejection device arranged on the Mars vehicle, initial speed is provided for the detection aircraft, and the Magnus rotor rotates to provide lift force;
step 2: in the climbing flight stage, the detection aircraft is quickly pulled to a certain height, and in the stage, the landing gear is retracted to reduce the air resistance, so that the parafoil is ensured to have enough height to open the parachute after entering the parachute opening cruise stage;
and step 3: in the parachute opening and cruising stage, after the detection aircraft climbs to a certain height, the parafoil is opened, the magnus rotor rotates and enters a cruising state, and information collection is carried out on the surface of the mars;
and 4, step 4: and in the gliding and landing stage, the energy supply system controls the magnus rotor to stop rotating, and the landing point is controlled by adjusting the parafoil to realize directional landing.
The Mars cruising detection method provided by the embodiment of the invention at least has the following beneficial effects:
1. the method for Mars cruise detection utilizes the fact that the flying height of a detection aircraft is low, and therefore high-precision remote sensing detection can be achieved only by carrying a detection load with low weight.
2. The moving speed of the detection aircraft in the air is much higher than that of a common mars vehicle, the comprehensive detection of one area can be completed in a short time, and the area with important phenomena can be revisited in a short time to be continuously monitored.
3. The detection of the detection aircraft is less affected by the terrain, and the detection aircraft can reach the area which cannot be reached by the mars vehicle and land to carry out in-situ detection, so that the planning of the detection task can be quite flexible.
4. The detection aircraft can also carry out advanced detection on the landform and the possible phenomenon in front of the mars train, and provides key data for scientists to determine which direction the mars train moves next.
According to the embodiment of the invention, in the step 2, in the parachute opening cruise phase, the attack angle of the detection aircraft is controlled to be 8-12 degrees.
According to some embodiments of the invention, in step 4, the angle of attack of the probe vehicle is controlled to be between 8 ° and 10 ° during the glide landing phase.
Additional aspects and advantages of the invention will be set forth in part in the description which follows and, in part, will be obvious from the description, or may be learned by practice of the invention.
Drawings
The above and/or additional aspects and advantages of the present invention will become apparent and readily appreciated from the following description of the embodiments, taken in conjunction with the accompanying drawings of which:
FIG. 1 is a schematic diagram illustrating a launch initiation phase of an aircraft according to an embodiment of the invention;
fig. 2 is a schematic diagram of the parachute cruise phase of the probe aircraft.
Reference numerals:
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to the same or similar elements or elements having the same or similar function throughout. The embodiments described below with reference to the accompanying drawings are illustrative only for the purpose of explaining the present invention, and are not to be construed as limiting the present invention.
In the description of the present invention, it should be understood that the orientation or positional relationship referred to in the description of the orientation, such as the upper, lower, front, rear, left, right, etc., is based on the orientation or positional relationship shown in the drawings, and is only for convenience of description and simplification of description, and does not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus, should not be construed as limiting the present invention.
In the description of the present invention, a plurality of means is one or more, a plurality of means is two or more, and greater than, less than, more than, etc. are understood as excluding the essential numbers, and greater than, less than, etc. are understood as including the essential numbers. If the first and second are described for the purpose of distinguishing technical features, they are not to be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated or implicitly indicating the precedence of the technical features indicated.
In the description of the present invention, unless otherwise explicitly limited, terms such as arrangement, installation, connection and the like should be understood in a broad sense, and those skilled in the art can reasonably determine the specific meanings of the above terms in the present invention in combination with the specific contents of the technical solutions.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
A probe vehicle according to an embodiment of the present invention is described below with reference to fig. 1 and 2.
As shown in fig. 1 and 2, a probe aircraft according to an embodiment of the present invention includes:
the wing 100 is in a wing shape, the wing 100 is provided with an energy supply system, and the wing 100 is provided with detection equipment;
the magnus rotor 200 is rotationally arranged at the front end of the wing 100, the magnus rotor 200 is arranged in parallel with the front end of the wing, and the energy supply system is in transmission connection with the magnus rotor 200 to drive the magnus rotor 200 to rotate;
the parafoil 300, the parafoil 300 is arranged on the upper surface of the wing 100, and the energy supply system is electrically connected with the parafoil 300 to supply energy to the parafoil 300.
The application of this application for surveying aircraft on the wing based on magnus effect installs magnus rotor 200 additional at wing 100 leading edge, and magnus rotor 200 is at the rotation in-process, can provide bigger lift for wing 100, and magnus rotor 200 can provide necessary lift demand for mars detector through measures such as changing the rotational speed under the low density low dynamic pressure condition of mars atmosphere.
The magnus rotor 200 is additionally arranged at the front end of the wing 100, and the device can provide lift force under certain incoming flow conditions. And compared with the traditional wing profile, the magnus rotor 200 can provide higher lift force at lower incoming flow speed, the lift force change is more stable, and the reliability is higher. For the magnus effect and the application thereof, reference may be made to patent application document cn201921199625.x, which is a smart mars train capable of taking off and landing vertically, and patent application document CN110435928A, which are detailed descriptions, and thus, detailed descriptions thereof are omitted.
Specifically, in the non-working stage, the detection aircraft is arranged in a detector cabin inside the mars vehicle and forms a structural relation with the mars vehicle, after receiving a working instruction, the mars vehicle opens the cabin door, and the ejection device ejects the detector to enter the working stage.
The leading edge of the wing 100 is additionally provided with a magnus rotor 200 with Li-CO2The battery pack is used as a power source for the rotor to operate, a speed measuring device is additionally arranged at the front edge of the wing 100, the rotating speed of the rotor is adjusted at any time through the measured real-time rotating speed, and the rotating speed ratio is guaranteed to be always kept at a predicted value. The rotational speed ratio here refers to the ratio of the circumferential rotational speed of the magnus rotor 200 to the incoming flow speed.
Meanwhile, a gliding parafoil is additionally arranged above the wing 100, so that on one hand, the severe mars weather can be avoided as required, the landing place of the paradise can be controlled, on the other hand, higher lift force is provided, and the bearing capacity of the paradise is improved.
Numerical simulation is carried out under the condition that the flight attack angle of the detection aircraft is 8 degrees under the environment of the earth and the mars, and the simulation result is shown in table 1:
TABLE 1 aerodynamic coefficients for combined airfoils and non-rotor airfoils under different atmospheric environments
It can be seen that in the earth environment, the advantage of the combined wing profile in aerodynamic performance is mainly reflected in the great improvement of the lift coefficient. In a Mars environment, the lift-drag ratio of a pure wing type aircraft is obviously reduced, vortex shedding occurs at the trailing edge of the wing, and the flight performance of a rotor-free wing on a Mars is greatly reduced, so that the application of the traditional fixed wing aircraft in the field of Mars detection is greatly limited.
Although the lift-drag ratio of the combined wing profile is lower than that of the combined wing profile in the earth environment due to the reduction of Mars atmospheric density, the reduction degree of the combined wing profile is far lower than that of the wing profile without the rotor, and the lift-drag ratio can be kept about twice of that of the wing profile without the rotor.
In some embodiments of the present application, the energy supply system includes a solar array that covers the outer surface of the airfoil 100 and a rechargeable lithium-carbon dioxide battery disposed inside the airfoil 100, the solar array being capable of recharging the lithium-carbon dioxide battery.
The atmospheric atmosphere on the Mars contains 96 percent of carbon dioxide, and the detection aircraft adopts Li-CO2The battery pack is provided with resources by high-concentration carbon dioxide on Mars, and is Li-CO2The ideal working environment of the battery pack. Mixing Li-CO2The battery pack is arranged in the detection aircraft and supplies power to the Magnus rotor 200, the parafoil 300, the landing gear, the detection device and the like. In some embodiments of the present application, the detector aircraft employs a solar cell array and Li-CO2The battery pack is integrated and jointly powered by MPPT energy transmission mode, Li-CO2The battery pack is directly hung on a bus, the voltage of the bus is kept between 2 and 4.5V, and energy is provided for the load of the whole detector. Single probe flight with Li-CO2The battery pack is mainly used for supplying power, and the solar battery array is used for charging the battery pack after the detection task is finished so as to be used for the next detection.
In some embodiments of the present application, the detection device is disposed inside the wing and is disposed at a middle front position of the wing 100, a lower surface of the wing 100 is of a transparent structure, and the detection device detects through the lower surface of the wing 100.
The middle front part of the lower surface of the wing 100 is made of a material with high transparency, so that the design has the advantages that the image information collection on the earth surface of the mars cannot be influenced, the visual angle is wider, the detection range is wider, the raising moment of the combined wing profile can be properly balanced while the aerodynamic effect of the wing profile is not influenced, and the stability of the wing profile is improved.
In some embodiments of the present application, the bottom of the wing 100 is provided with a landing gear 110. The landing gear 110 is arranged on the lower surface of the wing 100 to support the aircraft to realize fixation and sliding on a mars vehicle, realize stable landing on the surface of the mars in the sliding and landing stage, and retract the landing gear in the takeoff stage and the cruise stage to reduce air resistance.
In some embodiments of the present application, the magnus rotor 200 is rotated in a clockwise direction. Factors influencing the aerodynamic performance of the detection aircraft are many and include rotor speed ratio, wing-type parafoil spacing, incoming flow attack angle, atmospheric environmental parameters and the like. Changing the speed ratio significantly affects the aerodynamic performance and stability of the combined airfoil. When the rotation direction of the rotor is clockwise, the lift coefficient of the combined wing profile is higher, and the aerodynamic characteristics are more excellent.
A numerical simulation experiment is carried out, and three groups of rotation speed ratios are selected, namely a rotation speed ratio-4 rotor anticlockwise rotating, a rotation speed ratio 0 rotor not rotating and a rotation speed ratio 4 rotor clockwise rotating. The distance between the wing profile and the wing umbrella is twice of chord length of the wing profile, the incoming flow speed is 10m/s, the attack angle is 6 degrees, and the earth atmospheric environment is selected. The results are shown in table 2:
TABLE 2. influence of rotor rotation direction on aerodynamic characteristics of composite airfoils
Through the numerical simulation experiment, the maximum lift-drag ratio can be obtained when the magnus rotor 200 rotates clockwise.
In some embodiments of the present application, the rotational speed ratio of the circumferential rotational speed of the magnus rotor 200 to the incoming flow speed is 0-4. The speed of the incoming flow is measured by means of a velometer arranged on the wing, so that the rotational speed of the magnus rotor 200 is regulated and controlled, and the rotational speed ratio of the circumferential rotational speed of the magnus rotor 200 to the speed of the incoming flow is controlled to be 0-4.
The rotor rotation direction is determined to be clockwise. Four groups of rotation speed ratios, namely 0, 1, 2, 3 and 4, are selected to explore the influence of the rotor rotation speed ratio on the aerodynamic performance and the flow field characteristic of the combined airfoil profile. Mars atmospheric conditions are selected, the incoming flow speed is 10m/s, and the attack angle is 6 degrees. The simulation results are shown in table 3:
TABLE 3 influence of rotation speed ratio on aerodynamic characteristics of combined airfoils under Mars atmosphere
The optimum speed ratio of 4 is shown in Table 3. The probe vehicle can maintain the most efficient flight.
In some embodiments of the present application, the ratio of the length of the parachute line to the length of the chord length of the wing of the parafoil is 1-2.
Considering that the length of the parachute cord of the conventional parafoil is limited, and the length of the parachute cord limits the distance between the parachute 300 and the wing 100, from the practical application point of view, the combined wing type with the one-chord length distance and the two-chord length distance is selected, and the lift-drag performance and the stability performance of the combined wing type are discussed in a comparison mode. The simulated structures are shown in tables 4 and 5:
TABLE 4 aerodynamic coefficients of different pitches at an angle of attack of 6 °
TABLE 5 aerodynamic coefficients at different pitches at an angle of attack of 10 °
The angle of attack is 6 respectively for the combination wing section of contrasting respectively, when 10, along with the increase of interval, lift-drag ratio can slightly increase, and its pitch moment coefficient is less simultaneously, and stability is also better, nevertheless compares in the influence of angle of attack, and the interval is less to the aerodynamic parameter influence of combination wing section.
From the numerical simulation result, when the spacing between the wing and the parafoil is controlled to be twice the chord length, the pneumatic performance is optimal, the overlarge spacing does not meet the practical requirement, the parachute ropes are easy to break, the too small spacing can reduce the pneumatic performance of the combined wing profile, and therefore the parachute ropes with the twice chord length are the optimal rope length.
The invention also discloses a Mars cruising detection method, which comprises the following steps:
step 1: in the ejection starting stage, the detection aircraft is ejected and lifted off through an ejection device arranged on the mars vehicle, initial speed is provided for the detection aircraft, and the magnus rotor 200 rotates to provide lift force;
step 2: in the climbing flight stage, the detection aircraft is rapidly pulled to a certain height, and in the stage, the landing gear 110 is retracted to reduce the air resistance, so that the parafoil 300 has enough height to open the parachute after entering the parachute opening cruise stage;
and step 3: in the parachute opening and cruising stage, after the detected aircraft climbs to a certain height, the parafoil 300 is opened, the magnus rotor 200 rotates and enters a cruising state, and information is acquired on the surface of the mars;
and 4, step 4: in the gliding and landing stage, the energy supply system controls the magnus rotor 200 to stop rotating, and the landing point is controlled by adjusting the length of the parachute rope of the parafoil 300, so that the directional landing is realized.
Specifically, the electromagnetic catapult 500 is installed in the mars 400, the guide rail 410 is arranged in the mars 400, the detection aircraft is placed at the starting point of the guide rail 410, when catapult is performed, the power system transmits current to the coil on the track 410, and due to the electromagnetic induction effect, a magnetic field which continuously moves forwards and changes is generated along the track. The magnetic field interacts with a 7-rotor coil magnetic field on the reciprocating vehicle below the guide rail to generate thrust to push the reciprocating vehicle to move forwards and accelerate, so that the detection aircraft is driven to fly away from the mars vehicle in a bouncing manner, and the rapid take-off of the detector is realized.
When the detection aircraft flies to a sufficient height, the power supply system can supply power for the parafoil opening device for a short time, the parafoil is opened, the magnus rotor 200 at the front edge can be controlled to stop rotating for a short time at the moment, the lift force is reduced, the purpose is to enable the parafoil to reach a fully inflated state more quickly, and after the parafoil is fully opened, the control rope of the parafoil can be operated to control the flight direction of the detection aircraft, so that the detection aircraft enters a cruising working stage.
In the cruising process, the Li-CO2 battery pack is used as a power source in the working state, and the high-concentration carbon dioxide on the mars can provide resources for the battery pack, so that the battery pack is an ideal working environment for the Li-CO2 battery pack, and long-endurance flight of the detector is ensured; meanwhile, the solar sailboard is laid on the surface of the machine body, and the solar sailboard is used for charging the battery pack after single detection flight is finished so as to be used for next detection after recovery, so that the reuse rate of the detector is improved, and the service life is prolonged.
After cruising, the energy supply system stops supplying power to the magnus rotor 200, controls the magnus rotor 200 to stop rotating, and detects that the aircraft gradually falls down due to the reduction of the lift force. In the falling process, the falling point is controlled by adjusting the length of the parachute line of the parafoil 300, and directional falling is realized. Generally, there are two methods for directional steering of ram-type parafoils: the first pull-down edge; the second method is to close several air inlets on one side of the outer wing by using a control rope, so that the single-side outer wing of the parafoil is folded down.
First, the method for detecting the flying height of the detection aircraft relative to the satellite detection can greatly reduce the flying height, and therefore, the remote sensing detection with higher precision can be realized only by carrying a detection load with small weight.
Secondly, the moving speed of the detection aircraft in the air is much higher than that of a common mars vehicle, the comprehensive detection of one area can be completed in a short time, and the area with important phenomena can be revisited in a short time to be continuously monitored.
In addition, the detection aircraft adopts a cruise detection mode, so that the detection aircraft is less influenced by terrain, can reach an area which cannot be reached by the Mars vehicle and land at the area for in-situ detection, and therefore the planning of a detection task can be quite flexible.
In addition, the detection aircraft can also carry out advanced detection to the topography and the topography in front of the mars car and the phenomenon that probably takes place, provides key data for scientists to decide in which direction the mars car next step moves.
Some embodiments of the invention, in step 2, control the angle of attack of the probe aircraft to be between 8 ° and 12 ° during the parachute opening cruise phase.
The angle of attack, also called angle of attack, is the angle between the direction of the line connecting the leading edge and the trailing edge of the detected aircraft wing profile and the direction of the airflow.
Through numerical simulation experiments, five groups of attack angles of 8-12 degrees are selected, other variables are controlled to be consistent, the rotating speed ratio of a rotor is 4, the incoming flow speed is 10m/s, and the results are shown in a table 6:
TABLE 6 aerodynamic characteristics of 8-12 degree angle of attack range in parachute opening cruise phase
According to the comparison of numerical simulation results of the attack angles of 8-12 degrees, when the attack angles are 8-12 degrees, the lift-drag ratio is slightly reduced along with the increase of the attack angles, and the reduction speed is gradually increased along with the increase of the attack angles; on the other hand, from the stability analysis, the excessively high attack angle can cause the pitching moment coefficient to rise in the descending trend of the stability of the detection aircraft, so that the optimal attack angle is 8 degrees.
Some embodiments of the present invention, in step 4, control the angle of attack of the probe vehicle to be between 8 and 10 during the glide landing phase.
Through numerical simulation experiments, five groups of attack angles from 8 degrees to 12 degrees are selected, other variables are controlled to be consistent, the rotor stops rotating, the incoming flow speed is 10m/s, more detailed numerical simulation calculation is carried out on the detection aircraft in the glide landing stage, and the result is shown in table 7:
TABLE 7 aerodynamic characteristics of glide landing stage over an angle of attack range of 8-12 deg
The comparative analysis shows that the lift-drag ratio of the detection aircraft is obviously reduced along with the increase of the attack angle within the range of 8-12 degrees in the glide landing stage, and the lift-drag ratio is rapidly reduced when the attack angle exceeds 10 degrees. A significant drop in flight performance has occurred earlier than the stall during the parachute opening cruise phase because the rotor stall reduces the stall angle of attack of the fuselage, allowing the stall to occur at a lower angle of attack.
Stall angle of attack: if the angle of attack of the wing is large to a certain extent, the wing is equivalent to a flat plate erected in the airflow, and because the angle is too large, the airflow streamline bypassing the upper wing surface cannot be connected, separation can occur, and the airflow streamline flows backwards and downwards under the driving of the outer layer airflow and finally is rolled into a closed vortex, namely a separation vortex, so that the pressure in the rotating vortex is constant and is equal to the pressure of the airflow above the vortex. Therefore, the pressure difference between the upper wing surface and the lower wing surface is much smaller, and the lifting force of the wing is reduced compared with the original lifting force. To some extent, stall is formed, and the corresponding wing angle of attack is called the "stall angle of attack".
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an illustrative embodiment," "an example," "a specific example" or "some examples" or the like are intended to mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above do not necessarily refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.
While embodiments of the invention have been shown and described, it will be understood by those of ordinary skill in the art that: various changes, modifications, substitutions and alterations can be made to the embodiments without departing from the principles and spirit of the invention, the scope of which is defined by the claims and their equivalents.
Claims (9)
1. A probe vehicle, comprising:
the aircraft comprises a wing, an aircraft body and a control system, wherein the wing is in a wing shape and provided with an energy supply system, the wing is provided with detection equipment, the detection equipment is arranged inside the wing, and the bottom of the wing is provided with an undercarriage;
the magnus rotor is rotatably arranged at the front end of the wing and is arranged parallel to the front end of the wing, and the energy supply system is in transmission connection with the magnus rotor to drive the magnus rotor to rotate;
the parafoil is arranged on the upper surface of the wing, and the energy supply system is electrically connected with the parafoil to supply energy to the parafoil.
2. The probe vehicle of claim 1, wherein: the energy supply system comprises a solar cell array and a rechargeable lithium-carbon dioxide cell, the solar cell array covers the outer surface of the wing, the lithium-carbon dioxide cell is arranged inside the wing, and the solar cell array is connected with the lithium-carbon dioxide cell to charge the lithium-carbon dioxide cell.
3. The probe vehicle of claim 1, wherein: the detection device is arranged at the middle front position of the wing, the lower surface of the wing is of a transparent structure, and the detection device penetrates through the lower surface of the wing to perform detection.
4. The probe vehicle of claim 1, wherein: the magnus rotor rotates clockwise.
5. The probe vehicle of claim 1, wherein: the rotation speed ratio of the circumferential rotation speed of the Magnus rotor to the incoming flow speed is 0-4.
6. The probe vehicle of claim 1, wherein: the ratio of the length of the umbrella rope of the parafoil to the length of the chord length of the wing is 1-2.
7. A method for Mars cruise detection, characterized by comprising the following steps:
step 1: in the catapult starting stage, the detection aircraft in any one of claims 1 to 6 is catapulted and lifted off through a catapult device arranged on the mars carriage, and initial speed is provided for the detection aircraft, and at the moment, a magnus rotor rotates to provide lift force;
and 2, step: in the climbing flight stage, the detection aircraft is quickly pulled to a certain height, and in the stage, the landing gear is retracted to reduce the air resistance, so that the parafoil is ensured to have enough height to open the parachute after entering the parachute opening cruise stage;
and step 3: in the parachute opening and cruising stage, after the detection aircraft climbs to a certain height, the parafoil is opened, the magnus rotor rotates and enters a cruising state, and information collection is carried out on the surface of the mars;
and 4, step 4: and in the gliding and landing stage, the energy supply system controls the magnus rotor to stop rotating, and the landing point is controlled by adjusting the parafoil to realize directional landing.
8. The method of mars cruise detection according to claim 7, wherein: in the step 3, in the parachute opening cruise stage, the attack angle of the detection aircraft is controlled to be 8-12 degrees.
9. The method of mars cruise detection according to claim 7, wherein: in the glide landing stage, the attack angle of the detection aircraft is controlled to be 8-10 degrees.
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| CN114030644B (en) * | 2021-07-19 | 2024-02-09 | 中南大学 | Fixed wing aircraft lift-increasing device suitable for Mars atmosphere and aircraft |
| CN113911317A (en) * | 2021-11-08 | 2022-01-11 | 河南科技大学 | Wing and fixed wing aircraft |
| CN115979329B (en) * | 2022-05-19 | 2023-10-13 | 中国科学院国家空间科学中心 | Full-elevation expansion ball falling in-situ detection device and method for near space |
| CN115320827B (en) * | 2022-10-14 | 2023-01-31 | 中国航空工业集团公司沈阳空气动力研究所 | High-lift-drag-ratio airfoil profile with high subsonic speed and low Reynolds number flow |
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| JP4086384B2 (en) * | 1998-11-24 | 2008-05-14 | 富士重工業株式会社 | Aircraft automatic guidance system with parafoil and its navigation guidance device |
| US20030025038A1 (en) * | 2001-08-06 | 2003-02-06 | Lockheed Martin Corporation | Autonomous control of a parafoil recovery system for UAVs |
| US6860449B1 (en) * | 2002-07-16 | 2005-03-01 | Zhuo Chen | Hybrid flying wing |
| DE20216089U1 (en) * | 2002-10-19 | 2003-03-06 | Wolff Robert | Auxiliary aircraft rotor for e.g. airplanes, comprises rotary body parallel to wing for generating extra upthrust |
| US20110042521A1 (en) * | 2008-11-18 | 2011-02-24 | Sample Daniel S | Spacecraft Launch and Exploration System |
| DE102010055676A1 (en) * | 2010-12-22 | 2012-06-28 | Eads Deutschland Gmbh | hybrid rotor |
| US8448898B1 (en) * | 2012-04-30 | 2013-05-28 | Sunlight Photonics Inc. | Autonomous solar aircraft |
| CA2822665C (en) * | 2012-07-31 | 2018-07-17 | Gabor Devenyi | Aircraft wing having continuously rotating wing tips |
| CN103434640B (en) * | 2013-08-25 | 2015-07-15 | 西北工业大学 | High-altitude unmanned plane equipped with Magnus effect propulsion system |
| DE202016001030U1 (en) * | 2016-02-17 | 2016-03-04 | Alfons Malecha | wing |
| CN110077580A (en) * | 2019-04-28 | 2019-08-02 | 中国航空工业集团公司西安飞机设计研究所 | A kind of aircraft |
| CN110422341A (en) * | 2019-07-31 | 2019-11-08 | 南京航空航天大学 | A kind of Marsokhod kite balloon airship system and its working method for mars exploration |
| KR20190116156A (en) * | 2019-08-21 | 2019-10-14 | 서준영 | Self powered motor using magnus effect |
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