141results about How to "Improve aerodynamic characteristics" patented technology

To provide a cowl structure for a vehicle for improving the appearance of the vehicle without increasing the weight of the vehicle and also for enhancing the operational stability. In a cowl structure for a vehicle which includes a front cowl for covering a front portion of a vehicle body, a radiator is positioned in front of the cylinder heads of the engine. A radiator cover is provided for covering the radiator ranging from side surfaces to a lower end of the radiator. Bulging portions are provided for bulging from a front portion to a rear portion of the radiator cover and being formed in a direction to be rearwardly and upwardly oblique in a side view of the vehicle. The bulging portions are formed on a lower end portion of the radiator cover.

The invention provides a skin of a morphing wing, which is made from fiber-reinforced composite material (1). The skin is characterized in that: a shape memory alloy driver (2) is arranged between the two layers of the fiber-reinforced composite material (1); the wave trough of the upper layer fiber-reinforced composite material is internally filled with silica-based rubber (3) for meeting the requirements of smoothness, continuity and whole gas tightness on the surface of the wing; the fiber-reinforced composite material (1) and the shape memory alloy driver (2) are jointly pressed into to a corrugated structure under room temperature; the shape memory alloy driver (2) is in the shape of a high-temperature memory straight line; and the silica-based rubber (3) can be solidified under normal temperature. The invention also provides a drive method of the skin of the morphing wing. The skin of the morphing wing has the corrugated structure, thereby having high flexibility, large morphing amount and fast morphing response speed on the morphing direction, having high carrying capability on the perpendicular morphing direction, and having simple structure, light weight and reliable performance.

The invention provides a variable camber self-adaptive wing structure, the whole wing is divided into wing leading edge section, wing trailing edge section and each spanwise fixed wing structure section therebetween, wherein the The wing structure segment connected to the edge structure segment is fixed to the fuselage through the spar, and these structural segments are connected to each other by a double-fin structure. This double-fin structure allows the force between the wing structure segments to pass through the ribs. The transmission of the contact surface avoids the influence of the hinge gap of the wing deformation mechanism on the deformation, and improves the structural weight and complexity increased to meet the structural strength requirements of the wing. Use the control system of the aircraft to control and change the camber of the wing, so that the aircraft can obtain the optimal aerodynamic efficiency at different speeds and angles of attack. This adaptive wing with variable camber can greatly improve the lift. Enhancing the maneuverability of the aircraft and improving the stall characteristics can greatly improve the endurance time of the aircraft and the stability of the aircraft itself.

The invention provides a cylinder type launch folding wing unmanned aerial vehicle and an launch method thereof. The cylinder type launch folding wing unmanned aerial vehicle comprises a folding wingunmanned aerial vehicle (7) and a launch cylinder (6), wherein the folding wing unmanned aerial vehicle (7) is in a complete folding state and is arranged in the launch cylinder (6); when the launch cylinder (6) launches the folding wing unmanned aerial vehicle (7), the folding wing unmanned aerial vehicle (7) is ejected out of the launch cylinder (6), and climbs up to enter a complete unfolding loitering task state after a gradually unfolding varying process. The cylinder type launch folding wing unmanned aerial vehicle has the characteristics of simple structure, light weight and high engineering feasibility. Meanwhile, a Z-shaped folding elastic wing with the advantages of small land occupation in a folding state, big wing area in an unfolding state, large aspect ratio and the like is used to effectively improve the space use ratio and the aerodynamic performance of the cylinder type launch folding wing unmanned aerial vehicle in the loitering task state.

The invention discloses a near-space unmanned aircraft. Helium gas is inflated into an inflatable aircraft, solar wings are mounted at the front end of the inflatable aircraft, can rotate around flatter positions of rotary shafts at the roots of the wings by an angle ranging from -45 degrees to 80 degrees and can be locked at the three positions of a -45-degree angle, a 0-degree angle and a 80-degree angle. A plurality of propellers are mounted on the solar wings, two slantingly-rotatable trust propellers and two directional-trust auxiliary propellers are mounted on the inflatable aircraft, and solar cell panels are distributed on the surface of the solar wings and the inflatable aircraft. The inflatable aircraft is divided into a front section, a middle section and a rear section, wherein the rear section has a soft structure, a middle section has a semi-hard structure, and the front section is a task loading cabin. The near-space unmanned aircraft can overcome the defects of high requirements to taking-off and landing fields, insufficiency in effective load, poor safety in low-altitude flying and the like of an existing unmanned aircraft, and meanwhile, higher flight altitude can be realized, hovering posture and flight path control performances of the near-space unmanned aircraft are improved, absorption and storage of solar energy are improved and time for staying in space is guaranteed.

The invention provides a low-Reynolds-number wing section matched with a full-wing solar unmanned aerial vehicle. The low-Reynolds-number wing section is characterized in that the relative thickness of the wing section is 11%-13%, the maximum camber of the wind section is 2%-4%, and a chordwise position with the maximum camber is 25%-30%; the wing section has a single bent outline in a 70% chordwise range from a front edge. With a full-aircraft matching design, the checking of pneumatic characteristics of a pair of wing sections is finished by checking the corresponding pneumatic characteristics of the wing sections and applying the two wing sections to the full-wing solar unmanned aerial vehicle, and the low-Reynolds-number wing section shows good pneumatic characteristics and engineering enforceability and can meet application demands on low Reynolds number, high lift force and lift-to-drag ratio and large relative thickness of the full-wing solar unmanned aerial vehicle.

A method for aerodynamic robust optimization design of thick airfoil on that inner side of wind turbine blade According to the uncertainty of Reynolds number of incoming flow in the actual operation of wind turbine blades, Taking the maximum lift coefficient of airfoil and the probability statistics of stall characteristic parameters under random Reynolds number as the characteristic parameters ofairfoil affected by Reynolds number uncertainty, the robust optimization objective function of thick airfoil is formed by using the weight coefficient method and the aerodynamic lift characteristicsof airfoil at high angle of attack. A robust optimization design method for thick airfoil is proposed based on the geometrical characteristics of airfoil and genetic algorithm. The numerical prediction and evaluation of the optimization result of the concrete case show that the robust optimization design method of the large-thickness airfoil proposed by the invention can make the new airfoil obtain a high aerodynamic lift level in the range of a large angle of attack, and further improve the robustness of the aerodynamic characteristics of the airfoil with Reynolds number change, thereby realizing the design requirements.

The invention relates to the technical field of unmanned aerial vehicles, and particularly relates to an unmanned aerial vehicle with a tailed flying wing configuration, which has a plurality of working modes such as vertical take-off and landing / short-distance take-off and landing / conventional take-off and landing / low-speed forward flight / high-speed forward flight and the like and can realize coast-based / vessel-based multi-based configuration. The unmanned aerial vehicle comprises a lifting body type vehicle body of which the longitudinal section is of an airfoil shape, and main airfoils, upper airfoils and end airfoils which are arranged at both sides of the vehicle body and form smooth integration with the vehicle body, and a duct propeller propulsion device and a vertical fin which are arranged at the tail of the vehicle body, wherein the lower end of the vertical fin is fixedly connected with a duct structure of the duct propeller propulsion device arranged at the tail of the vehicle body. According to the invention, the tailed flying wing configuration integrated with jointed wings is adopted, so that the unmanned aerial vehicle has various advantages such as excellent aerodynamic characteristics and high effective load factor of the conventional flying wing configuration, and also effectively overcomes the inherent defects of the conventional flying wing configuration in aspects of stability and operability.

The invention belongs to the technical field of cross of bionics and aerodynamics, and discloses a novel feather-wing bionic ventilation flapping wing system and a multi-vortex interference mechanismanalysis method, so as to derive and describe a motion equation set and capture the vortex motion tracks. The method is characterized by analyzing the influence rules of the dynamic aerodynamic characteristics, the flow rules, the dynamic vortex structures and the vortex interference characteristics of the ventilation flapping wing system. The invention provides a multi-body model design of the novel feather wing bionic ventilation flapping wing system, and provides an aerodynamic shape and a flapping scheme, so that the design thought of the micro flapping-wing aircraft is expanded, and the aerodynamic characteristics are improved, the influence rules of the parameter changes of different fluttering frequencies, fluttering angles, fluttering inclination angles, Reynolds numbers and the like on the lift resistance characteristics, the thrust characteristics, the flow rules, the dynamic vortex structures and the like of the ventilation flapping wing layout are concluded and summarized,and the dynamic aerodynamic force between the moving winglet and the main flapping wing and a multi-vortex interference mechanism of the dynamic aerodynamic force are summarized.

The invention provides a complex-hinge transformable trailing edge mechanism based on planar linkage closed-loop units. The complex-hinge transformable trailing edge mechanism comprises at least three closed-loop linkage units and a drive mechanism. Each closed-loop linkage unit is a planar linkage structure based on a 6R closed-loop unit and has one freedom degree. The closed-loop linkage units include a tail-end unit, at least one middle unit and a starting-end unit which are parallelly arranged, and the middle units are arranged between the starting-end unit and the tail-end unit. Compared with the prior art, the complex-hinge transformable trailing edge mechanism has the advantages that the basic units of the complex-hinge transformable trailing edge mechanism are independent 6R closed-loop units, the units influence on another, and the output of the previous unit is the input of the next unit; the whole mechanism contains few linkages and is lightweight; the freedom degree of the whole mechanism is 1, the mechanism can be driven by only one drive motor, the weight of the mechanism is reduced, system control difficulty is lowered, and the gradual bending effect of a transformable wing is good due to the fact that multiple basic unit modules are used.

The invention provides a bionic stratospheric airship, and belongs to the technical field of aircraft design. The shape of a hull of the stratospheric airship is obtained according to the shape setting out of a portuguese man-of-war. The hull is internally provided with more than one main airbag, more than one auxiliary airbag and a heat-regulating airbag. The main airbags are filled with gas withthe density lower than air, the main airbags are disposed on an upper layer in the hull, and first gas valves are arranged on the upper parts of the main airbags. The auxiliary airbags are filled with the air, the auxiliary airbags are disposed on a lower layer in the hull, and second gas valves and blowers are arranged at the bottoms of the auxiliary airbags. The heat-adjusting airbag is arranged in a middle layer in the hull, and the heat-regulating airbag is internally provided with a thermal cycle device. The heat-regulating airbag is filled with a gas-liquid reversible control working medium. Under the working conditions of super-thermal and over-pressure, the working medium is converted from a gaseous state to a liquid state by the thermal cycle device, and the working medium is converted from the liquid state to the gaseous state by the thermal cycle device under the ultra-cold working condition.

To provide a cowl structure for a vehicle for improving the appearance of the vehicle without increasing the weight of the vehicle and also for enhancing the operational stability. In a cowl structure for a vehicle which includes a front cowl for covering a front portion of a vehicle body, a radiator is positioned in front of the cylinder heads of the engine. A radiator cover is provided for covering the radiator ranging from side surfaces to a lower end of the radiator. Bulging portions are provided for bulging from a front portion to a rear portion of the radiator cover and being formed in a direction to be rearwardly and upwardly oblique in a side view of the vehicle. The bulging portions are formed on a lower end portion of the radiator cover.

The invention discloses a solar united wing unmanned aerial vehicle and a control method thereof. The united wing unmanned aerial vehicle comprises a tail supporting rod, wings and pod assemblies, thewings comprise front wings, back rings and a vertical fin, and a left front wing, a wing body and a right front wing are integrated to form the front wings; the left front wing and the right front wing are symmetrically arranged at the two sides of the wing body, the wing body is fixedly connected with the tail supporting rod, and a left back wing and a right back wing which are symmetrically arranged are integrated to form the back wings; the back wings are located above the tail supporting rod, the connecting joint of the left back wing and right back wing is connected to the tail portion of the tail supporting rod through the vertical fin, the outer side ends of the left back wing and right back wing are in overlapping connection to the middles of the left front wing and right front wing respectively through streamline vertical end plates to form united wings, and four pod assemblies are arranged and respectively located at the lower portions of the front wings and back wings. Thesolar united wing unmanned aerial vehicle can solve the problems that an existing unmanned aerial vehicle with the quite high aspect ratio is poor in structure intensity and difficult in design, is only controlled by an electromotor, and cannot cooperate with a steering engine to achieve comprehensive control.

A drive and guidance apparatus for a trailing-edge / landing flap,-arranged on an aircraft mainplane, the apparatus comprising a six-element guide chain having seven rotating joints and one or more shafts; a first coupling element mounted via a crank connected to the mainplane wherein the crank can rotate, the coupling is connected to the flap and to a second coupling element wherein the first coupling can rotate, the second coupling mounted on the mainplane wherein the second coupling can rotate. The second coupling element is connected to the flap via an oscillating support. The crank is arranged wherein the rotation direction of the crank for extension of the flap corresponds to the rotation direction of the crank for the trailing-edge flap of the starboard wing of the mainplane, with the rotation direction of the crank for the trailing-edge flap of the starboard wing being counterclockwise when viewed from the aircraft fuselage.

The invention discloses a power-operated tail-sitting type mixed layout vertical take-off and landing aircraft. The power-operated tail-sitting type mixed layout vertical take-off and landing aircraft is composed of a fuselage, airfoils, motors, propellers and landing gears; the fuselage axis coincides with the OX axis of a body axial system, the projections of the fuselage axis and the OX axis in an XOY plane of the body axial system are distributed in an X shape; each airfoil on the fuselage is divided into several sections, sweepback angles and dihedral angles of all the sections are different from one another, and excellent pneumatic performance and maneuvering performance are achieved through positively-curved airfoil profiles and negatively-curved airfoil profiles; four sets of propeller-motor power systems are installed on the four airfoils correspondingly, and the distances between the positions where the propeller-motor power systems are located and the OX axis of the body axial system are the same; and four power devices take off in an X-shaped quad-rotor mode in the vertical take-off and landing processes and complete conversion operation to enter a cruising state through different tensile forces of the motors or by being matched with maneuvering surfaces, and required maneuvering is completed through tensile force changing of the motors in the whole process. The power devices of the aircraft are simple, the control mode is reliable, propeller slipstreams can be effectively utilized, and the aircraft is suitable for serving as flying platforms of a tail-sitting type vertical take-off and landing unmanned aerial vehicle.

The invention relates to a method for controlling the thermal effluents of an aircraft including an airframe (110) and at least one propulsion unit (112), wherein said at least one propulsion unit (112) includes a turbomachine (116) supplied with fuel via a fuel supply circuit (122) extending from a tank (124) provided at the airframe (110), said airframe (110) including at least one thermal effluent source (134), characterised in that it comprises scattering at least partially, at the level of at least one propulsion unit (112), the thermal effluents generated at the airframe (110) by using the fuel used for supplying the turbomachine (116) as a heat-carrier fluid.

To effectively perform lightning protection in a blade for wind power generation and to prevent the blade from being damaged when arresting lightning. The lightning protection structure of the blade for wind power generation includes a conductive lightning receptor 1 attached to a part of the blade for wind power generation, and a ceramic member 10 interposed between at least surface-layer portions of the lightning receptor 1 and the blade 5. Therefore, an electric field is likely to concentrate at an interface between the ceramic member 10 having excellent heat resistance and the lightning receptor 1, so that it is possible to effectively prevent the blade 5 from being damaged due to a concentration of the electric field at the interface of the blade 5 attached with the lightning receptor 1 when arresting the lightning.

The invention relates to a miniature ornithopter with a wing twisting function and multiple flight motion functions and belongs to the field of miniature ornithopters. A flapping torsional mechanism, an empennage, a fuselage direct-current motor, an empennage supporting rod and an electronic control module are installed on an ornithopter frame body. Wings are connected to torsional panels. An adjusting and steering rotor is connected to a fan motor. An empennage transmission mechanism is connected to the empennage supporting rod. The empennage supporting rod and the empennage are fixedly connected together by an empennage supporting rod fixing structure through screws. The miniature ornithopter has the beneficial effects that the structure is novel, the design structure is simplified, and miniaturization is easy to achieve. The ornithopter can generate larger pushing-up force compared with an aircraft only flapping up and down under the condition that the flapping frequencies are the same, the flight performance of the miniature ornithopter is greatly improved, the ornithopter flies flexibly, the aerodynamic characteristics are improved, the flight efficiency is high, the size is small, the weight is small, practicality is higher, and more flight functions can be achieved.

An end configuration for a bat includes in one embodiment an end cap that has a base that is shaped to fit within the open end of the bat thereby to attach the end cap to the bat. The end cap also includes a convexly contoured end surface that is exposed when the end cap is attached to the bat. The end surface is roughened to reduce the aerodynamic drag when the bat is swung, thereby to increase the momentum of the bat for a given amount of force applied to swing the bat.

The invention discloses a designing method of an efficient low-load wing section special for a horizontal shaft wind turbine blade and the wing section obtained on the basis of the method. The designing method is based on an anti-design method, and the designed wing section is efficient, low-load, good in variable working conditions, and smooth in stalling. The efficient low-load wing section has the advantages that a designing goal of traditional wing sections is changed, and a pneumatic characteristic required by the wing section is provided; the wind section has high maximum lift-drag ratio, high designed lift coefficient and low maximum lift coefficient, and low lift coefficient decreasing under the condition front edge roughness is achieved; the anti-design method is utilized to design the wing section; and corresponding pneumatic parameter designing indexes are provided aiming at wing section applied to the outer sides, with thickness not more than 25%, of wind turbine blades.

The invention discloses a rotary aircraft wing folding and unfolding mechanism and a folding method thereof, which mainly solve the problem existing in the prior art that the conventional folding mechanism occupies the internal space of the wing, the moving mechanism is not compact and the wing surface is uneven due to the single rotating hinge. The rotary aircraft wing folding and unfolding mechanism comprises an inner wing, an outer wing, a driving gear, and a folding actuator for driving the driving gear; an inner wing link mechanism and an outer wing link mechanism are respectively articulated with the opposite surfaces of the inner wing and the outer wing; a driven gear A and a driven gear B are respectively fixedly connected to the inner wing link mechanism and the outer wing link mechanism; the driving gear, the driven gear A and the driven gear B are all incomplete gears; and at different stages, the driving gear is engaged with the driven gear A or the driven gear B. Through the solution, the invention achieves the objectives of even wing surface and compact structure, having high practical value and popularization value.

The invention discloses a slice omnidirectional broadband surface conformal antenna, which is formed by an antenna (1), a coaxial feeding joint (3), an A resistance (4) and a B resistance (5), wherein the structure of the A resistance (4) is same with the structure of the B resistance (5), the sensitive end of a probe (35) in the coaxial feeding joint (3) is welded on the feeding end (1a) of the bottom end of the antenna (1), the coaxial feeding joint (3) is mounted on the lower plate surface (2b) of a mounting plate (2), the A resistance (4) is mounted between the bottom of a right antenna wing (12) of the antenna (1) and the upper plate surface (2a) of the mounting plate (2), and the B resistance (5) is mounted between the bottom of a left antenna wing (13) of the antenna (1) and the upper plate surface (2a) of the mounting plate (2). The slice omnidirectional broadband surface conformal antenna simultaneously adopts bottom-end two-side resistance loading and a bottom-end conical gradation structures which lead the band width of the slice omnidirectional broadband surface conformal antenna to be widened. Compared with the bowed antennas which are used at present, the surface conformal antenna with a slice structure can reduce the negative influence to platform pneumatic property, the installation position is flexible, and a place which is convenient for antennas to play efficiency is chosen to mount.

The invention relates to a flow guide type air barrier. An air barrier support is composed of upright columns, an upper beam and a lower beam, and forms an air barrier unit. The dip angle between the axis of each upright column of the air barrier and a plumb line ranges from 1 degrees to 45 degrees. Multiple rows of transverse air barrier bars and multiple columns of vertical air barrier bars are distributed in the air barrier support and are connected smoothly on the windward side to form a whole air barrier face. The vertical air barrier bars are arranged continuously at equal intervals between the upper beam and the lower beam, and dip angles ranging from 1 degree to 45 degrees are formed between the vertical air barrier bars and the plumb line. The upper beam and the lower beam of the air barrier can be integrated or combined. Flow guide air barrier bars are additionally arranged on the windward side of the whole air barrier face. By means of the air barrier, a part of air flow can bypass an air barrier plate, so that the flow guide type air barrier has a wind shielding function as a common air barrier and also has an air flow guide function. When the air barrier is applied to a noise barrier, wind resistance of the noise barrier can be improved. By means of the air barrier, the pneumatic characteristics o f a bridge structure can be improved, and wind resistance of the bridge structure is improved.

The invention discloses a deflection control device based on the head of an aircraft. The deflection control device is composed of a deflection mechanism, the head of the aircraft, an intelligent skin, an intelligent framework and an aircraft body; the deflection mechanism is installed between the head of the aircraft and the aircraft body; two ends of the intelligent framework are connected with the aircraft body and the head of the aircraft, respectively; and the intelligent skin is fixed on the intelligent framework. The deflection control device is used for controlling by taking the head of the aircraft as a control plane of the aircraft, a conventional canard is replaced, and negative aerodynamics and thermo-mechanical effects caused by the fact that the head is interfered with the airflow of the canard can be avoided; because the pressure increment of the front end is generated due to mutual deflection of the head and the canard, non-uniform change is generated in local circumstances along the range of the canard; simultaneously, the control plane is at the front part of the aircraft, so that the operation efficiency is higher; the head of the aircraft directly pulls shock wave under supersonic velocity so as to obtain higher aerodynamic force; and the deflection control device is simple in aerodynamic arrangement structure and better in aerodynamic characteristics and the aircraft resistance is low.

The invention discloses a horizontal scanning conformal array antenna. The horizontal scanning conformal array antenna comprises a mounting chassis and multiple antenna units which are uniformly distributed on concentric circles taking a center of the chassis as a circle center, wherein each antenna unit is connected with a coaxial feeding cable, a lower surface of an antenna shield is mounted on the mounting chassis, a phase control electric scanning device is connected with the antenna units through the coaxial feeding cables, and wave beam scanning in all directions can be realized through controlling phase of each antenna unit. The conformal array antenna is advantaged in that the structure is simple, the antenna can be conformal with an unmanned plane or a missile platform and can be hidden in a carrier sheath, good pneumatic characteristics and good electric index of a carrier are maintained, a radiation field of the antenna can realize uniform high gain scanning in a horizontal 360-DEG scope, and cooperation networking demands of the unmanned plane or a missile communication system can be satisfied.

The application of the invention provides a wind turbine blade and a wind turbine. The wind turbine blade includes a blade body and at least one flap flexibly mounted at the trailing edge of the blade body, wherein the flap is provided with two panels and an elastic component; the two panels are oppositely fixed on the two sides of the elastic component and rotate by taking the elastic component as a shaft, so as to form a meshed clamping structure at one end of the elastic component and forms a folding dovetail structure at the other end of the elastic component. The flap is meshed with the trailing edge of the blade body through the clamping structure, the clamping structure is fastened with the trailing edge of the blade body through the thrust applied on the inner side of the dovetail structure, of the elastic component, so that the aerodynamic characteristics of the blade body can be improved and further chattering of the blade body is reduced.

The invention discloses a wind turbine blade wing section group which comprises a first wing section, a second wing section and a third wing section. The outlines of the first wing section, the second wing section and the third wing section are formed by smoothly connecting non-dimensional two-dimensional coordinates, the non-dimensional two-dimensional coordinates are obtained by using horizontal coordinates and longitudinal coordinates of all points on the pressure faces and suction force faces of the first wing section, the second wing section and the third wing section to divide the chord lengths of the wing sections. The relative thicknesses of the first wing section, the second wing section and the third wing section are respectively 23%, 27% and 32%, and the maximum camber positions are respectively located at the positions of the chord lengths of 67.6%, 74.1% and 64.4% of the front edges of the wing sections. The demands for thick wing sections of large wind turbine blades applied to a low-wind-speed wind field are met, the aerodynamic characteristics of the thick wing sections are effectively improved, the maximum lift coefficient and maximum lift-drag ratio of the thick wing sections are improved, and front edge roughness sensibility of the wing sections is reduced. Compared with existing thick wing sections, the structural performance of the blades is obviously improved, and static load under the limiting condition is reduced, so that the blades are lengthened, wind diffusion area is increased, and more wind energy is captured.

A pulsating heat pipe active thermal protection structure coupled with ablative resistant materials comprises a thermal grooming device and ablative resistant materials, wherein the ablative resistantmaterials are coated on the periphery of the thermal grooming device and comprise an inner layer and an outer layer; the inner layer is a flexible foam, and the outer layer is a high thermal conductivity carbon / carbon material; the thermal grooming device comprises a cooling panel, a pulsating heat pipe group and a phase change cold storage box; the cooling panel is bonded with the ablative resistant materials; the pulsating heat pipe group is welded with the cooling panel; the phase change cold storage box is welded with the cooling panel and fixed with the end frame of the aircraft; and thepulsating heat pipe extends into the phase change cold storage box. The pulsating heat pipe active thermal protection structure coupled with ablative resistant materials combine ablation-resistant material with the thermal grooming device. By means of the active thermal protection structure of the sharp lead edge of the aircraft, ablation is reduced or no ablation occur, and the aerodynamic shapeof the aircraft is changed little, and the aerodynamic characteristic of the aircraft with high lift-to-drag ratio is maintained, and the problem of material failure caused by the local high heat ofthe aircraft is solved, and the extreme distribution of temperature is avoided, and the long-time heat insulation is realized.

The invention discloses an all-directional wideband antenna which has the same shape with an installation surface structure; the all-directional wideband antenna comprises a coaxial feed joint, a cylindrical back chamber (1), a glass fiber reinforced plastic cover (2), a metal upper cover (3), a metal lower cover (7), a sleeve (8), three groups of short circuit columns and three supporting frames; three groups of short circuit columns are arranged between the metal upper cover (3) and the metal lower cover (7); the three supporting frames are arranged on the internal wall of the sleeve (8) which is welded inside the cylindrical back chamber (1) and is positioned at the jointing of a conical bottom surface (101) and a transition section (103); the glass fiber reinforced plastic cover (2) is arranged on a flange platform (104) of the cylindrical back chamber (1); the frontal disc (35) of the metal upper cover (3) is attached to the glass fiber reinforced plastic cover (2); the frontal disc (35) of the metal upper cover (3) is attached to the glass fiber reinforced plastic cover (2), thus leading the all-directional wideband antenna to be internally embedded into the installation plane of a movable platform so as to form the structural conformality.