511 results about "Aerospace technology" patented technology

The invention discloses a nonlinear self-adaption control method of near-space hypersonic vehicle (NHV), which belongs to a flight control method in the technical field of aerospace. The control method mainly comprises three control law parts: a nominal nonlinear generalized predictive control law (NGPC), a B-spline recursive functional linkage network (BRFLN) self-adaption control law, and a robust control law of gain self-adaption adjustment. The invention integrates the simplicity of the NGPC method and the effectiveness of dynamic uncertainty of the BRFLN learning, targets the immeasurable dynamic uncertainty and fast interference of an attitude system in the flight of the NHV, brings favorable learning effect, and realizes the nonlinear accurate control to the attitude angle.

The invention discloses a non-cooperative target satellite acquisition device and an acquisition method, and belongs to the technical field of aerospace. The non-cooperative target satellite acquisition device includes an outer shell, two four-link capturing pawl mechanisms, four horizontal locking mechanisms, a screw rod and a controller; the two four-link capturing pawl mechanisms are arranged in the outer shell face to face, and the lower parts thereof are fixedly connected with an installing bottom plate; the lower end of the screw rod crosses through the installing bottom plate, and the lower end thereof is connected with a driving mechanism; four horizontal locking mechanisms are respectively located at outside of the two four-link capturing pawl mechanisms, and every two horizontal locking mechanisms are installed on the upper part of the outer shell; both sides of the upper part of the four-link capturing pawl mechanisms are respectively equipped with pressure lever mechanisms; when the four-link capturing pawl mechanisms are descended, the pressure lever mechanisms compress the horizontal locking mechanism to horizontally glide towards inside; the top part of the outer shell is equipped with a photoelectric sensor, and the photoelectric sensor and the driving mechanism are electrically connected with the controller. The non-cooperative target satellite acquisition device is widely applied to the acquisition of various space crafts, and the non-cooperative target satellite acquisition device is wide in application range, big in tolerance, quick in response speed, high in reliability, and others.

The invention relates to an unmanned helicopter performance testing platform, belongs to the field of aerospace technologies and aims at solving the problems that the fixing structure is unstable, the balance performance in the movement process is poor and a testing table is damaged in the prior art. The unmanned helicopter performance testing platform comprises a platform body assembly and a flight simulator, wherein an unmanned helicopter is fixedly arranged on an upper platform of a six-degree-of-freedom moving platform, and lifting force data generated during the flight of the unmanned helicopter is transmitted to a computer through a lifting force sensor A which is fixedly arranged on the lower surface of a lower platform of the six-degree-of-freedom moving platform and a lifting force sensor B which is fixedly arranged on the lower surface of an iron plate module, so as to realize performance testing. The unmanned helicopter performance testing platform has the beneficial effects that the action of the unmanned helicopter is provided by virtue of the six-degree-of-freedom moving platform and a positioning sleeve can prevent the helicopter from flying away from the testing table when the flight control of the helicopter fails or an experiment fails; since the lower surface of the lower platform of the six-degree-of-freedom moving platform is seated on the upper surface of a shock absorbing plate, the balance of the unmanned helicopter is kept; the shock absorbing plate, a large compression spring and a small compression spring also play a certain role in buffering.

The invention relates to a POD (proper orthogonal decomposition) and surrogate model based order reduction method for hypersonic aerodynamic thermal models and belongs to the technical field of aerospace. A hypersonic aircraft aerodynamic thermal environment is predicted and obtained by the aid of a POD and surrogate model method, nonlinear characteristics of real gas effect, turbulence viscosity and the like of high-precision numerical calculation are reserved, the hypersonic aircraft aerodynamic thermal environment can be provided for design of the hypersonic aerocraft by the aid of model order reduction method which is high in precision and efficiency, related thermal boundary conditions are provided for aerodynamic thermal elastic design, the thermal environment is provided for thermal protection design of the hypersonic aerocraft, design efficiency can be greatly improved, design circle is shortened, and design cost is reduced.

The invention relates to a minimum base attitude disturbance track planning method for a redundant space manipulator, and belongs to the technical field of aerospace. The method comprises the steps of (1), designing a three-dimensional model of the redundant space manipulator; (2), establishing a kinematic model of the redundant space manipulator; (3) establishing a kinematic equation of the redundant space manipulator by employing a generalized Jacobian matrix; (4), parameterizing joints by employing a 5-order sine polynomial function; (5), establishing a cost function equation under a minimum constraint condition of base attitude disturbance; (6), providing an improved hybrid particle swarm optimization; and (7), carrying out optimization solution to the cost function by employing the hybrid particle swarm optimization, thereby obtaining planning tracks under the constraint condition. According to the method, the track planning problem of the redundant space manipulator under the minimum constraint condition of the base attitude disturbance is solved, and the planning tracks solved by the improved hybrid particle swarm optimization is high in precision and good in effect and is steady.

The invention discloses a shock-free wrapping tape type unlocking separation mechanism, belongs to the technical field of aerospace, and aims at solving the problems of high production cost, high shock response, long working time, small load and pollution. Connecting bolts of two shock-free connection unlocking assemblies are connected with a wrapping tape type separation assembly through wrapping tape gaskets and nuts; a memory alloy barrel is heated to be recovered to the original shape; a disconnection shaft is pulled and broken into two parts; an actuation spring pushes a guide barrel to move; connecting bolts are released; and a left wrapping tape and a right wrapping tape are pushed away by a separation spring and separated from an upper supporting seat and a lower supporting seat, so that the upper supporting seat is separated from the lower supporting seat. According to the invention, the three-stage release process is achieved, so that the bearing capacity can be amplified step by step on one hand and the shock response can be minimized through step-by-step release at the same time; a tension spring and the actuation spring can effectively absorb shock generated by brittle failures of the bolts; and the shock-free connection unlocking assemblies can bear a larger clamping force and the wrapping tape type unlocking separation mechanism can bear a large load.

The invention relates to an interplanetary transfer orbit design method, in particular to a transfer orbit design method from a periodical orbit which is near dynamic balance points of a three-body system to a small celestial body and belongs to the technical field of aerospace. Firstly, based on the periodical orbit which is near the dynamic balance points of the three-body system where a detector is located, an initial state of the detector x0=[ r0, v0] is assumed. Secondly, a speed increment delta v is exerted along a tangential direction so that the detector sets off from the periodical orbit near the dynamic balance points. Based on a multiple-body model, relations between the detector and a distance df of the detector and the target small celestial body, a flying period tf and the tangential direction speed increment delta v are built up when the detector flies across the target small celestial body is built. Through an optimizing algorithm, an initial value of the transfer orbit is obtained when the fight distance df is minimum. Finally, a secondary-level differential correction method is used based on the initial value to obtain orbit parameters which can meet two point boundary values. The interplanetary transfer orbit design method can achieve rapid design of flight from the periodical orbit which is near the dynamic balance points of the three-body system to the small celestial body, calculation amount is small and efficiency is high.

The invention discloses a balance point Halo orbit phasing orbit transfer method taking the time constraint into consideration, relates to a Halo orbit phasing orbit transfer method based on an earth-moon three-body dynamic model and belongs to the technical field of aerospace. According to the method, a dynamic equation is established under a restrictive three-body model formed by the earth, the moon and the star, and a Halo orbit near the point L2 is generated under an earth-moon rotation system; the Halo orbit initial phase position of a detector and the phase position difference needing to be changed are determined, and the optimal fuel phasing orbit meeting the phase position constraint condition and the transfer time constraint condition is obtained through an optimization algorithm with the initial anchoring time and transfer time as optimization variables; and the time difference delta t, the task Halo orbit or the upper limit tmax of the transfer time are adjusted according to the task which the detector needs to fulfill, and the orbit shadow detection avoiding task or the spatial intersection detection task of the detector on the Halo orbit is fulfilled. The method can obtain the optimal fuel phasing orbit meeting the phase position constraint condition and the transfer time constraint condition, and has the advantages of good convergence, high flexibility and the like.

A method for producing a fibre composite component, in particular for aerospace, the method comprising the following steps: forming a mould core from a material comprising cork by a moulding tool to establish an outer geometry of said mould core; arranging the so formed mould core adjacent to an at least partly hardened stiffening element on a base element of said composite component to be produced for the shaping of at least one moulded portion of said fibre composite component to be produced; and multistage exposure of at least said moulded portion to heat and / or pressure to produce said fibre composite component; a corresponding mould core and a corresponding fibre composite component.

The invention discloses a preparation method of a POSS (Polyhedral Oligomeric Silsesquioxane) modified polysiloxane rubber composite material. The preparation method is characterized in that the POSS containing single / double reactive functional groups is introduced into a polysiloxane molecular chain containing a plurality of reactive functional groups by virtue of grafting or blocking in a chemical copolymerization manner, and then a POSS modified polysiloxane oligomer with the POSS as a polysiloxane hanging chain and the POSS as a polysiloxane block chain, next, 0.1-100 parts of POSS modified polysiloxane oligomer is evenly dispersed in 100 parts of polysiloxane raw rubber by virtue of physical blending, and finally, 0-50 parts of white carbon black and 0.2-10 parts of a curing agent are added to the mixture for cross-linking and solidification and then the POSS modified polysiloxane rubber composite material is obtained. The POSS modified polysiloxane rubber composite material has excellent high and low temperature resistance, mechanical properties and dielectric property, and can be applied to the technical fields of automobiles, electronics, building materials as well as national defense and aerospace.

The invention relates to a method for monitoring plume contamination of an attitude control engine of a spacecraft, in particular to a method for monitoring plume contamination of an attitude control engine of a spacecraft by adopting a quartz-crystal microbalance, belonging to the technical field of aviation. In a monitoring device for the plume contamination of the attitude control engine of the spacecraft, a set of quartz-crystal microbalance monitor is arranged on the sliding rail of the existing space electric propulsion ground simulation testing equipment, and two sets of quartz-crystal microbalance monitors are arranged behind an engine nozzle on the surface of a vacuum chamber. In the method, an attitude engine is adopted to ignite so as to ensure that the engine jets heat flow gas, and contaminants are generated when the gas is condensed; and the frequency and temperature of the quartz-crystal microbalance are monitored online in situ and saved. The method improves the sensitivity of plume contamination monitoring of the attitude control engine in China reaching 1.10*10-9 to 4.42*10-9, and the testing process is stable and reliable and has good repeatability, so the method is suitable for large-scale tests.

The invention discloses a space floating object capturing device in the field of aerospace technology. The space floating object capturing device comprises a floating platform with a propelling device, an auxiliary adjusting device and a capturing device; the floating platform is connected with an aerospace carrier through a flexible mechanism; the flexible mechanism comprises multiple sets of wire roller systems fixed on the surface of the aerospace carrier, multiple sets of wire rolling mechanisms arranged in the aerospace carrier or the floating platform, and wire sets which have two ends respectively connected with a dragging motor and a floating platform and pass through the wire roller systems; According to the space floating object capturing device, the floating platform provided with the propelling device is controlled by the plurality of wires, the capturing device is carried by the platform, the floating platform is driven by the jet propelling device to be close to a target, and the posture and angle of the platform are adjusted by controlling the lengths of wires connected with the floating platform and the auxiliary adjusting device. The space floating object capturing device has the advantages of simple structure, small size, easiness in implementation of control, being capable of mounting in carious aerospace devices and capable of recycling for use and the like.

The invention discloses a large-scale four-dimensional flight path dynamics prediction method and device, and belongs to the technical field of aerospace. The device comprises an information obtainingmodule, a data pre-processing module, a track clustering module, a dynamic data access module, an identification and prediction module, and a result feedback module. The information obtaining modulesamples aviation data to constitute flight path sequences. The data pre-processing module removes noise points on each flight path sequence and re-samples the flight path sequences without the noise points. The track clustering module is used for clustering the flight path sequences. The dynamic data access module accesses real-time data, to obtain a four-dimensional vector at equal time intervalsto form flight path sequences. The identification and prediction module recognizes the abnormal flight mode and carries out trajectory prediction. The result feedback module reports the real-time flight path prediction and abnormal recognition conditions to the air traffic management department to guide flight operation in the airspace. Airspace smooth and safe operation is achieved, and great importance is achieved.

The invention discloses a minor planet detection optimal multi-impulse transfer method of an interplanetary multibody system, and belongs to the technical field of aerospace engineering. The method includes steps of firstly, building a detector kinetic equation under a geological core rotating system, namely, a high-precision kinetic model; selecting a target minor planet according to task restriction and providing an initial state of the detector under the geological core rotating coordinate system; applying a disturbance method to obtain a detector single-impulse minor planet flyover track;applying an optimization algorithm to obtain a single-impulse minor planet flyover transfer track with the minimum flyover distance; based on a main vector theory, performing a multi-impulse transfertrack design, and acquiring an optimal multi-impulse transfer track which meets the main vector condition; bringing the acquired optimal multi-impulse transfer track to the high-precision kinetic model to progressively induce; amending the transfer track by a multi-grade parallel differential amending method, and realizing the accurate track transfer of minor planet detection under a multi-celestial-body strong disturbance nonlinear environment. The minor planet detection optimal multi-impulse transfer method has the advantages of small speed increment, good applicability, and good astringency.

The invention discloses a new preparing method of MnCoNi film material through wetting-chemical method, which is characterized by the following: enlarging area at lower temperature; stabilizing the heat-sensitive material of MnCoNi film; breaking the limit of traditional ceramic element to satisfy the request of aviation and astronautic technological element.

The invention provides a Halo orbit on-orbit keeping method considering amplitude constraint, and belongs to the aerospace technology field. The method comprises the steps that by establishing a kinetic equation under a limited three-body model composed of two main celestial bodies and a detector, the position of a balance point of a three-body system composed of the two celestial bodies and the detector is determined; a Halo orbit, near the balance point, in the three-body system composed of the big and small celestial bodies and the detector is determined; a differential correction algorithm for Halo orbit keeping is designed according to the disturbance variable; a Halo orbit on-orbit keeping strategy is designed according to the differential correction algorithm; in a real ephemeris environment, measurement errors and execution errors are considered, the speed increment of each time of orbit correction is obtained according to the orbit keeping strategy through the differential correction algorithm, orbit correction control is conducted according to the speed increments. Accordingly, Halo orbit on-orbit keeping considering amplitude constraint can be achieved, and consumption of fuel needed in orbit keeping can be reduced as much as possible.

The invention discloses a multi-step typed ion jet device and a method, belonging to the fields of aerospace technology and microelectronic technique; the device comprises a first accelerating electrode, a second accelerating electrode, and an end hollow electrode; the first accelerating electrode and the second accelerating electrode are alternatively arranged so as to form a multi-step electrode structure; an insulation layer is arranged between every two adjacent first accelerating electrode and second accelerating electrode so as to lead the first accelerating electrode and the second accelerating electrode to be insulated with each other; the method utilizes the electric field convergence effect of polarized electrode array tip part, reinforces the electric field strength of region adjacent to the electrode, thus ionizing the gas molecular of adjacent area; the multi-step ion jet device structure is adopted and voltage is applied on the structure, thus forming ion jet. The device and the method can improve the ionization efficiency, reduce the working voltage, and work under the environment of multi-gas-plasma.

The invention belongs to the field of aerospace technologies, and relates to a multi-directional flexible supporting structure of an oversized-aperture optical reflector. The multi-directional flexible supporting structure of the oversized-aperture optical reflector aims to solve the problems that in the prior art, it is difficult to simultaneously meet high surface precision requirements and high assembly process level requirements of a reflector under the working conditions of gravity release and varying environment temperature. The multi-directional flexible supporting structure of the oversized-aperture optical reflector comprises a large-caliber reflector body, taper sleeves and flexible joints, wherein three supporting holes are evenly distributed in the 120-degree circumference of the back of the large-caliber reflector body, the taper of each supporting hole is the same as that of the corresponding taper sleeve, each supporting hole and the corresponding taper sleeve are matched and are bonded and fixed through glue, each flexible joint is fixed to the corresponding taper sleeve through screws and pins, and the upper portion of each flexible joint is connected with a connecting boss on a back supporting plate of the large-caliber reflector body through screws. The three flexible joints are evenly distributed on the back of the large-caliber reflector body, each flexible joint is provided with four transverse grooves and two longitudinal grooves, and it is ensured that the reflector will not have obvious deformation caused by variation of the external environment and the imaging quality of the reflector will not be affected.

The invention provides a small reflector ultrahigh power thermal stability supporting structure, and belongs to the technical field of the aerospace technology. The supporting structure solves the problems that in the prior art, a reflector is low in intrinsic frequency and poor in external disturbance resistance, a reflection mirror surface is large in shape change, the requirement for the assembly technique level is high, interchangeability is poor and mass production is difficult to achieve. The supporting structure comprises a mirror chamber and a flexible joint, the flexible joint and the reflector are in glue joint and fixed concentrically, the flexible joint is fixed in the mirror chamber and comprises a cylindrical barrel and multiple flexible single arms which are uniformly distributed on the outer side of the cylindrical barrel in the circumference direction and each flexible single arm is a double-layer flexible hinge. Under the action of temperature loads, in the process that thermal deformation stress deformation energy of the lens barrel is transmitted to the reflector, the thermal deformation stress deformation energy is absorbed when passing a flexible unload groove of the flexible joint, uniform deformation of the reflector is guaranteed, at the moment, surface precision slightly changes, the parts are fixedly connected, the requirement for the assembly technique is not high, and the supporting structure is good in interchangeability and high in external dynamic interference resistance.

The present invention relates to a minimum base disturbance trajectory planning method for a space robot, in particular to a minimum base disturbance trajectory planning method for a space robot based on an improved chaotic particle swarm algorithm, which belongs to the field of aerospace technology; the steps included in the present invention To: (1) establish the kinematic equation of the 6-DOF space robot; (2) parameterize the joints with the 7th-order sine polynomial function; (3) establish the equation of the fitness function, including the constraints of the position of the base Disturbance, attitude disturbance, joint velocity and angular velocity constraints; (3) In order not to fall into local optimum and quickly find the optimal solution, the chaotic particle swarm optimization algorithm is improved; (4) The improved chaotic particle swarm optimization algorithm is used to solve the unknown Parameters, the optimal trajectory of the space robot can be obtained in the case of the minimum disturbance of the base posture. The invention focuses on solving the problem of trajectory planning for a 6-degree-of-freedom free-floating space robot, and uses the improved chaotic particle swarm algorithm to find out the optimal trajectory of the space robot when the disturbance of the base seat posture is minimal, and the effect is good.

The invention discloses a lunar finite thrust multi-capture orbit realization method and belongs to the field of aerospace technology. According to the specific realization method, a prober finite thrust kinetic equation is established under a lunar inertial system; a single lunar capture maneuver is decomposed into multiple capture maneuvers, and parameters of each capture maneuver are optimized;and velocity pulses needed by the first maneuver and the remaining maneuvers are converted into an optimal finite thrust capture orbit, an initial work phase, a thrust direction and work time of an engine are optimized, optimal finite thrust capture is realized, and a target orbit is entered. Through the lunar finite thrust capture orbit realization method, the needed velocity increment is small,capture efficiency is high, the finite thrust capture orbit meeting constraint can be obtained, and convergence is good.

The invention discloses a six-degree-of-freedom active and passive dynamic vibration-absorbing device, which comprises a bracket and six single degree-of-freedom dynamic vibration absorbers, wherein the six single degree-of-freedom dynamic vibration absorbers are in pairwise symmetrical arrangement, and are fixedly connected with three vertical directions of the bracket via bolts respectively; and the vibration absorbers are connected with a controlled structure in two ways, namely, the vibration absorbers are connected internally and externally with the controlled structure. The bracket is a hollow metal framework connected by six crosses into a cube; each single degree-of-freedom dynamic vibration absorber comprises an end cover, a shell, a coil sleeve, a motor rotor, a diaphragm spring, an M4 nut and an M3 bolt. The six-degree-of-freedom active and passive dynamic vibration-absorbing device is installed on the surface of the controlled structure, has six-degree-of-freedom vibration suppression capability and active and passive control modes, is flexible and easy to control, fast in response speed and good in reliability, and is adapted to the complex vibration suppression working condition with multiple degree-of-freedom. The device has higher practical value and wide application prospect in the technical field of aviation and aerospace.

The invention relates to a mode switch valve of a variable-cycle engine adjustable mechanism and belongs to the technical field of highly efficient gas turbine engines in the aerospace techniques. The mode switch valve is used for changing the work state of the engine to adapt to various flight conditions and is arranged at the inlet of an outer duct of the engine. The mode switch valve comprises an outer case, a valve sheet, a seal ring, a connecting rod, a synchronous ring, a hydraulic actuating cylinder, a front gas seal case, a rear gas seal case and a seal expansion ring. The mode switch valve is matched with a front duct ejector and a rear duct ejector of the variable-cycle engine adjustable mechanism to adjust, the switch of two cycle modes of the engine which are turbofan and turbojet is realized, so that the engine has an optimal performance in a wide work scope.

The invention relates to flame-retardant and drippage-free ceramization silicon rubber and a preparation method thereof. The silicon rubber is prepared from components in parts by weight as follows: 100 parts of organic silicon rubber compounds, 1 part to 10 parts of a vulcanizing agent, 100 parts to 150 parts of ceramization powder and 0 to 100 parts of other filler or auxiliaries, wherein the ceramization powder is prepared from raw materials in parts by weight as follows: 1 part to 120 parts of glass powder, 1 part to 20 parts of muscovite, 1 part to 200 parts of magnesium hydroxide, 1 part to 200 parts of aluminum hydroxide, 0.1 parts to 50 parts of boron compounds and 0.1 parts to 20 parts of phosphono-containing polymer compounds. The ceramization silicon rubber prepared with the method has advantages of good flame retardance and self-extinguishment performance; a ceramic layer formed after firing has good compactness and insulativity, external force resistance and the like. The preparation process is simple, raw materials are cheap, and accordingly the silicon rubber is expected to be applied to fields such as wires and cables, the electronic and electrical technology, aerospace engineering and the like and has quite good economic benefits and popularization value.

The invention discloses a determining method for a structural stiffness index of a high-aspect-ratio wing, and belongs to the technical field of aerospace engineering. The determining method mainly includes the following steps: step 1, analyzing and determining structural stiffness design requirements of the wing; step 2, creating a structural stiffness analysis model of the wing; step 3, analyzing structural stiffness distribution, meeting the design requirements, of the wing; step 4, calculating a structural stiffness distribution index curve of the wing by a stiffness simulation analysis platform. According to conventional wing structural design methods, the stiffness design requirements are regarded as starting points, preliminary structural design is firstly performed according to design experiences, stiffness check is then performed, and the structural design is modified repeatedly according to check results to meet the stiffness design requirements. According to the determining method, for passivity and defects of the conventional wing structural design processes, the stiffness distribution index curve is provided before starting of the structural design to guide the wing structural design, so that iteration modification processes are reduced, and scientificity and design efficiency of the wing structural design are improved.

The invention relates to an in-situ monitoring method for outgasing contamination of nonmetallic materials on an optical surface of a spacecraft, belonging to the technical field of aviation. In the method, a quartz wafer is adopted to monitor contamination level on the optical sensing surface of the spacecraft in situ, and is arranged in a vacuum chamber; a nonmetallic material is arranged in the visual range of a quartz-crystal microbalance to ensure that the material outgases in the vacuum condition; a spectrophotometer and a vacuum system are adjusted, and the nonmetallic material is heated to lead contaminants to escape; frequency and temperature of the quartz-crystal microbalance are monitored online in situ and stored; and the system is closed. The in-situ monitoring method for the outgasing contamination of the nonmetallic materials on the optical surface of the spacecraft has high sensitivity reaching 1.10*10-9 to 4.42*1-9, and the testing process is stable and reliable, and has good repeatability. The method is suitable for large-scale tests.

The invention relates to an orbital element based interstellar flight launch opportunity searching method, in particular relates to an interstellar flight detector launch opportunity searching and selecting method, and belongs to the field of aerospace technology. The method comprises the following steps of: firstly determining an optimal dual-impulse rendezvous opportunity of a target celestial body and a launch celestial body; establishing the corresponding relationship between ephemeris of the launch celestial body and the mean anomaly in launch; calculating the rendezvous time of the launch celestial body and the target celestial body by the ephemeris of the launch celestial body; acquiring the state of the target celestial body in rendezvous through the ephemeris of the target celestial body, and converting the state into the orbital element, so as to obtain the mean anomaly of the target celestial body corresponding to the preset rendezvous time; and if the mean anomaly is matched with the optimal mean anomaly, judging that the launch opportunity is the optimal launch opportunity of the area, otherwise, readjusting the date corresponding to the optimal launch mean anomaly. The orbital element based interstellar flight launch opportunity searching method provided by the invention has the advantages and effects that the interstellar flight optimal launch opportunity can be searched rapidly with less calculation quantity and high efficiency, meanwhile, the cycle property of the launch opportunity can be accurately recognized and the like.

The invention belongs to the technical field of aerospace, and relates to a parameter resembling method for aerodynamic thermal environment experimental simulation conditions. According to the requirement of equivalence of recovery enthalpy and cold wall heatflow in an aerodynamic thermal environment experiment, a method of guaranteeing ground test parameters being similar to flight parameters without recovery enthalpy equivalence, and the problems at the present stage that experimental simulation method depends on recovery enthalpy equivalence and accurate simulation of the real aerodynamic heat environment is difficult are solved. A guarantee for heat transfer and ablation experimental study of thermal shielding / insulation material in the high temperature aerodynamic thermal environmentis provided. The parameter resembling method has the beneficial effects that experimental conditions are not affected by recovery enthalpy value; most test equipment can simulate the high enthalpy aerodynamic thermal environment; a new way of parameter adjustment for high enthalpy wind tunnel testing is provided; and the experimental simulation conditions are broadened.

The invention discloses a low-orbit constellation disposition method based on assistance of martian atmosphere, and relates to a martian constellation disposition method under the martian atmosphere and a gravity system thereof. The low-orbit constellation disposition method belongs to the field of aeronautical and space technologies. According to the low-orbit constellation disposition method, a speed pulse which is applied for entering the atmosphere from an original orbit and a speed pulse that is applied for making an aircraft enter an objective orbit are solved through optimizing a control rate which satisfies an aerodynamic force requirement. A detector releases the carried aircraft at an apareon position and makes the aircraft enter the atmosphere through applying the speed pulse which is required for making the aircraft enter the atmosphere from the original orbit. In the atmosphere, aerodynamic-assisted orbit transferring is performed through optimizing the preset control rate. Furthermore the orbit of the aircraft is fixed to the objective orbit through applying the speed pulse which is required for making the aircraft enter the objective orbit. A plurality of constellation aircrafts are disposed to the objective tracks which correspond with the constellation aircrafts in a one-to-one manner, thereby realizing disposition of a whole constellation. The low-orbit constellation disposition method has advantages of low energy consumed in the disposition process, no rigid requirement for the objective orbit, and flexible disposition process.