40 results about "Space probe" patented technology

The invention relates to a satellite-borne electronic hardware design guarantee method and system. The method comprises the following steps: determining a satellite-borne hardware guarantee level (SHL) of satellite-borne electronic hardware; integrating a fault mode injection FM I into a design process of satellite-borne electronic hardware, and adjusting the execution depth of the FM I based on the SHL level; and providing a process cutting engine, and realizing adaptive process cutting based on a DO-254 standard according to the SHL level and in combination with the execution depth of the FM I. The system comprises an SHL evaluation module; an FM I management module; a space environment effect knowledge base; and a process cutting engine. According to the method, the problem that the DO-254 standard is insufficient in adaptability in the aerospace field is solved, the reliability design and verification efficiency of satellite-borne electronic hardware is remarkably improved, and the method is suitable for aerospace tasks such as various satellites and deep space probes.

The invention relates to the technical field of deep space exploration, in particular to a cryogenic propellant on-orbit low-loss storage system and method for a deep space probe. After the low-temperature propellant is injected into the storage tank through the injection pipe, along with continuous accumulation of external leaked heat, the first pressure sensor detects that the pressure in the storage tank is continuously increased, and when the numerical value of the first pressure sensor exceeds the preset pressure, the cooling assembly is started to cool the low-temperature propellant in the storage tank, so that the external leaked heat is consumed; before the engine needs to supply the low-temperature propellant, the low-temperature propellant in the storage tank is in a supercritical state, and the storage temperature is increased in a high-pressure storage mode, so that the temperature difference between the storage tank and the outside and the heat leakage amount are reduced, and the requirement for the refrigerating capacity of the cooling assembly is reduced; when the low-temperature propellant is conveyed, due to the fact that the low-temperature propellant is in a supercritical state and free of gas-liquid phase separation, the interior of the whole storage box is in a homogeneous state, the stable conveying flow of the low-temperature propellant can be guaranteed, and stable output of an engine is also facilitated.

The invention discloses a self-adaptive thermal management device of a shape memory Janus array structure, a preparation method of the self-adaptive thermal management device and application of the self-adaptive thermal management device in a deep space probe, and belongs to the technical field of self-adaptive thermal management. The problem that the heat dissipating capacity of a traditional radiation radiator with the single heat radiation characteristic cannot be adjusted is solved. The temperature response shape memory Janus array is adopted as a driver of the self-adaptive thermal management device, and switching of different thermal management modes under the condition of no external energy consumption is achieved. When the temperature is lower than the phase change temperature, the shape memory Janus array is in a pressed-down state, the low-emissivity coating coated on one side of the comb teeth is exposed, and the heat loss of infrared radiation is effectively inhibited by utilizing lower infrared emissivity. When the temperature is higher than the phase change temperature, the shape memory Janus array is in an inclined state, and the high-emissivity coating coated on the other side of the comb teeth is exposed, so that energy increase caused by solar radiation is prevented, and effective heat release is promoted.

[Problem] To provide novel technology. [Solution] One aspect of the present invention provides a thermionic cathode. The thermionic cathode includes an electron source and a plurality of grids. The electron source is configured to emit electrons when heated. The plurality of grids have through-holes that allow electrons emitted from the electron source to pass in a predetermined direction. At least one of the grids is configured to reduce the movement speed of the electrons.

This invention discloses an electrically driven unlocking and releasing device for a lunar space probe, relating to the field of high-altitude sensor deployment technology. The invention includes a base column, a housing fixedly connected to the top of the base column, a dome-shaped cover fixedly connected to the top of the housing, a release device at the top of the base column, and a stabilizing device on the inner wall of the housing. The stabilizing device includes a placement plate, the outer wall of which is fixedly connected to the inner wall of the housing. A bidirectional motor is fixedly connected to the bottom of the placement plate, and a threaded rod is fixedly connected to one end of the output shaft of the bidirectional motor. An unlocking disc is fixedly connected to the outer wall of the threaded rod, and an insert plate is fixedly connected to the top of the release cover. Through the stabilizing device, this invention increases the contact area between the unlocking disc and the insert plate by coordinating the bidirectional motor, threaded rod, unlocking disc, insert plate, insertion hole, locking disc, and locking hole, thereby improving the stability of the locking mechanism.

1. The name of the design product: Launch Vehicle (Chirei 6). 2. The use of the design product: used for space transportation, sending the payload such as artificial earth satellites, manned spacecraft, space stations or space probes into the predetermined space orbit. 3. The design points of the design product: the combination of shape, pattern and color. 4. The picture or photo that best shows the design points: perspective view. 5. The design claimed contains color.

The invention discloses a deep space probe landing point selection method based on topographic feature quantification, and the method comprises the steps: constructing a multi-target topographic feature optimization model through analyzing the parameters of a landing region, such as the topographic slope, the surface roughness, the obstacle distribution and the regional area, and achieving the automatic selection of a safe landing point. The method comprises the following specific steps: firstly, detecting spatial distribution of potential obstacles such as rocks and meteorite craters, and secondly, quantitatively analyzing topographic feature parameters such as topographic roughness, gradient change rate and topographic relief; and finally, an optimal landing position is determined under multiple terrain constraint conditions by using an optimization search algorithm, so that the landing safety and task success rate of the detector are effectively improved. According to the method, the safe landing point of the detector is accurately obtained by innovatively applying terrain feature quantification and a search algorithm, so that effective detection of a safe position in a landing area is efficiently realized, and safe landing of the detector is practically guaranteed.

The present invention relates to the field of deep space exploration technology, and in particular to a low-loss on-orbit storage system and method for cryogenic propellants used in deep space probes. After the cryogenic propellant is filled into the tank through a filling pipe, as external heat leakage continues to accumulate, the first pressure sensor detects that the pressure in the tank continues to rise. When the value of the first pressure sensor exceeds a preset pressure, the cooling component is activated to cool the cryogenic propellant in the tank, thereby absorbing the external heat leakage. Before the engine needs to supply the cryogenic propellant, the cryogenic propellant in the tank is in a supercritical state, and the storage temperature is increased by a high-pressure storage method, thereby reducing the temperature difference between the tank and the outside world and the amount of heat leakage, and thus the cooling capacity demand of the cooling component is also reduced. When the cryogenic propellant is transported, since it is in a supercritical state, there is no distinction between gas and liquid phases, and the entire tank is in a homogeneous state, which can ensure the stability of the cryogenic propellant delivery flow and is also conducive to the stable output of the engine.

The present invention provides a reusable airbag buffer landing mechanism and a deep space probe. The landing mechanism is characterized by comprising a main base (1), an airbag (2), support legs (3), and an inflation device (4). The main base (1) is the main board of the aircraft. The support legs (3) are fixedly installed on the main base (1). The airbag (2) is installed at the lower end of the support legs (3). The inflation device (4) is installed on the support legs (3). The airbag (2) has the function of reusable inflation and deflation. Before each landing, the airbag (2) is quickly inflated and deployed, and after landing, the airbag (2) gradually exhausts air. The present invention uses a reusable airbag to achieve repeated buffer landings of the probe, and has the advantages of reliable clamping and release, simple operation, convenient installation, good processing technology and economy. Furthermore, the present invention adopts a linkage mechanism combined with a buffer spring to assist in buffering, so as to enhance the stability of the airbag at the moment of landing and assist in buffer landing.

The invention discloses a lightweight space satellite detection positioning method, a hardware architecture and a space probe, and relates to the technical field of deep space exploration. The method comprises the following steps: receiving and preprocessing a space environment image; constructing a lightweight convolutional neural network model, wherein the lightweight convolutional neural network model comprises a multi-scale convolutional feature extraction and fusion module and a decoupling multi-scale detection head module; processing the received space environment image through a multi-scale convolution feature extraction and fusion module, and generating a feature map fusing multi-scale semantics and space details; and performing parallel classification prediction, foreground prediction and bounding box regression on the fused feature map through a decoupling multi-scale detection head module, and outputting a satellite target detection result. According to the method, semantic information and spatial details of target features are enhanced through multi-scale convolution feature extraction and fusion, the detection effect of a small target is remarkably improved, and the model can be deployed on a satellite-borne FPGA chip with limited resources through deep pruning and quantitative design.

The invention provides an observability analysis method and device for deep space exploration optical autonomous navigation, and the method comprises the following steps: constructing an orbit dynamics model and an optical observation model according to the measurement information obtained by a deep space probe; constructing an optical autonomous navigation system model according to the orbit dynamics model and the optical observation model; observability parameters of the deep space exploration optical autonomous navigation system are given according to the optical autonomous navigation system model; constructing an observability characterization model of the autonomous navigation system; carrying out dimensionality reduction characterization on observability of the deep space exploration optical autonomous navigation system; therefore, an observability criterion is obtained; and realizing the observability autonomous determination of the deep space exploration optical autonomous navigation system according to the observability criterion. By analyzing key factors of complete representation of the observability of the system, judgment conditions of the observability of the system are deduced, and the observability quantitative expression method which is universal, simple and easy to autonomously realize on a satellite is provided.

The invention belongs to the field of remote sensing satellite system engineering, and relates to a ground large closed-loop test system and method for a space exploration load working mode. A new space exploration load on-orbit working mode is designed according to the characteristics of an observation target, and the interior of the load is designed into a local sky area mode and a window tracking mode; the whole satellite load observation mode is designed into a natural rendezvous monitoring mode, a fixed-point staring monitoring mode, a scanning imaging mode and a target tracking mode; and aiming at a detection load mode, a dynamic star simulator is adopted to simulate an imaging sky area, and a simulated motion fragment target is added into the dynamic star simulator, so that simulation of a real imaging scene is realized. Meanwhile, in order to ensure that the load can successfully match and verify the local star catalogue and the sky area, a normal window tracking mode is switched to, so that the pointing direction corresponding to the sky area simulated by the dynamic star simulator is consistent with the pointing direction of the control broadcast, the on-orbit real working condition and time sequence can be accurately simulated, and the sufficiency of the ground test is improved.

The invention discloses a minimum number management system for deep space probe resource limitation. The minimum number management system comprises a processing module and a power distribution module which are located in a reentry capsule, wherein the power distribution module is responsible for supplying power to electronic equipment in the cabin and completing primary power supply combination, primary power distribution and control of a heater safety switch and a grounding switch; the power distribution module receives primary power input in the detectors or between the detectors, outputs primary power distribution to the inward processing module, and outputs primary power distribution and instruction power to other terminals in the detectors. Meanwhile, a safety switch of a heater control circuit of the processing module is controlled, and local state telemetering is externally output for inter-device state monitoring; the processing module receives the primary power supply input of the power distribution module and realizes secondary power supply conversion and power distribution; bus instruction input among detectors or in the detectors and external analog quantity, temperature quantity and passive switching quantity input are received, control instructions are output after logic processing, and data management and autonomous management are achieved at the same time.

[Problem] To provide novel features. [Solution] One aspect of the present invention sets forth a Hall thruster provided with an anode, a vaporization chamber, and a passage formation part. The anode is configured to supply a propellant to a predetermined position in order to generate a propulsive force. The vaporization chamber is configured to vaporize the propellant on the inside thereof. The passage formation part forms a first passage, a second passage, and a third passage. The first passage is configured to guide the propellant to the inside of the vaporization chamber. The second passage is configured to guide, of the propellant having been guided to the vaporization chamber, a predetermined amount of the propellant to the anode. The third passage is configured to guide, of the propellant having been guided to the vaporization chamber, the propellant other than the predetermined amount to the outside of the Hall thruster.

To provide a space probe which achieves weight saving and low cost in addition to improving reliability of a sample transport device.SOLUTION: A sample transport device of a space probe comprises: a link mechanism; a carrier which is connected with the link mechanism; and an actuator which drives the link mechanism. When A is a distance from a first support shaft to a first connection shaft, B is a distance from the first support shaft to a second support shaft, C is a distance from the second support shaft to a second connection shaft, and D is a distance from the first connection shaft to the second connection shaft, the following expression (1) is satisfied. The link mechanism is arranged, when the carrier is at an intermediate position of a movement path between a lower position and an upper position, so that the first support shaft, the second support shaft, and the second connection shaft are arrayed on the same line when viewed from an axial direction of the first support shaft. A+B<C+D ... (1)SELECTED DRAWING: Figure 3

The invention discloses a deep space probe lander power supply device suitable for Jinxing, Mars and South China, relates to the technical field of deep space exploration, and comprises a parachute which is jointly made of a fabric solar energy supply device, a canopy, a structural parachute cord, a conductive parachute cord and the like. The auxiliary power supply device is composed of an air bag, a parachute cord tractor and a conductive parachute cord winder. The device is characterized in that a fabric type solar energy supply device is adopted to replace a traditional canopy material, the photovoltaic power generation function is integrated on the parachute, meanwhile, an auxiliary power supply device is assisted to achieve the effects of electric connection or unfolding of the parachute on the surface of an extraterrestrial celestial body, and the huge surface area of the parachute serves as the effective solar power generation area; therefore, the power generation efficiency of the deep space detector in a low-illumination-intensity extraterrestrial environment is remarkably improved, and the defect that a traditional deep space detector is deficient in electric energy is overcome.

The present invention discloses a method for detecting mislocking of downlink signals of a deep space probe, which specifically includes the following steps: Step 1, calculate the transmission delay t of the downlink signal of the probe sy , before the time of T0 + t sy , if the ground locks the downlink signal, mislocking of the downlink signal is considered; otherwise, go to Step 2; Step 2, check the locking situation of the telemetry subcarrier. If the carrier is locked and the locking time of the subcarrier is less than the telemetry transmission time τ of 4 frames, it is preliminarily determined that the downlink signal is correctly locked; otherwise, the possibility of carrier signal mislocking is considered, and go to Step 3; Step 3, compare the difference between the measured and theoretical Doppler values. If the absolute value of this difference is greater than the limit value K f , mislocking of the signal is considered; otherwise, go to Step 4; Step 4, compare the difference between the actual signal-to-noise ratio and the theoretical signal-to-noise ratio of the downlink signal. If the absolute value of this difference is greater than the limit value K M , mislocking of the signal is considered; otherwise, return to the first step to re-enter the detection process. Using the present invention can effectively detect whether the deep space TT&C system mislocks the downlink signal of the probe.

To provide a propulsion device that vibrates a circular diaphragm by an external force, generates a difference in amplitude between two regions divided from the center of the diaphragm, and exploits the resulting disparity in vibrational energy.SOLUTION: Provided is a propulsion generation device that includes a vibration unit composed of a rotor 1 fixed with a weight 2 and a rotor bearing 5, and that linearly connects a motor shaft 7 fitted with a bevel gear 8 to a joint 6, with the bevel gears 8 of a pair of similarly configured vibration units engaging at an acute angle and arranged within a semicircular region of a circular vibration plate 4. The vibration plate 4 is secured to a vertical shaft 3 coupled to an intermediate portion of an outer frame 9 equipped with an adjustment screw 13, and a support frame 10 having a rotation shaft 11 orthogonal to the outer frame 9 is connected to a bottom plate 12 sharing the same rotation shaft 11. By synchronously rotating the weights 2 of the rotors 1, the device generates propulsion through differences in vibration energy.SELECTED DRAWING: Figure 1

This invention discloses a hierarchical matching pose estimation method for optical navigation landmarks in deep space probes, belonging to the field of autonomous navigation technology for deep space exploration. The implementation method is as follows: A probe pose dynamics model and an optical observation model are established to obtain the probe's initial position, attitude, and camera parameters, used to predict the probe's state and the set of visible landmarks at the observation time; a navigation landmark observability index function is constructed based on the Fisher information matrix; a hierarchical landmark selection and matching strategy is constructed, considering landmark visibility constraints, and selecting a primary landmark set, a secondary landmark set, and a candidate landmark set from the visible landmark set at the observation time based on the probe's trajectory and the set of visible landmarks at the observation time; sequential matching is performed during online matching to suppress the decrease in navigation information caused by landmark matching loss; and step-by-step matching is used during probe navigation to achieve rapid landmark identification and high-precision pose estimation.

This propulsion device generates a propulsive force by a combustion reaction, and includes: a combustion chamber, and a metal fuel supply device. The combustion chamber forms a combustion space inside of which it is possible to generate a combustion reaction. The metal fuel supply device is configured to supply a metal fuel to the combustion space. The supplied metal fuel is configured to extend in the combustion space, cause a combustion reaction, and vibrate during the combustion reaction as a result of theoperation of at least a part of the propulsion device.

This invention discloses a low-gravity simulation servo control method and apparatus for deep space exploration landing experiments. The method includes: acquiring the structural parameters of the low-gravity simulation test platform and establishing a spatial structural model of the parallel cable drive system based on the structural parameters; establishing a force-displacement characteristic model of the parallel cable drive system based on the spatial structural model, and obtaining mapping relationship data between platform displacement data and cable tension data based on the force-displacement characteristic model; acquiring the motion trajectory data of the deep space probe during the landing process, determining the target servo trajectory of the servo platform based on the motion trajectory data, and calculating the target cable tension data required to achieve the target servo trajectory based on the mapping relationship data; and controlling the drive device in the parallel cable drive system to adjust the tension of each cable based on the target cable tension data to drive the servo platform to move according to the target servo trajectory. This invention solves the technical problem of poor servo platform control performance in the prior art.

The invention relates to a space probe (2) with optimized thermal management for operation in extreme environments as encountered on the moon, the space probe comprising a chassis (4) to receive probe components, four wheels (10) attached to the chassis and a mobile body (14) assembled on the chassis, the mobile body comprising a wheel shield (16) movable between an open position in which the wheels are exposed and a closed position in which the wheels are covered, and an access hatch (18) movable between an open position in which at least some components of the probe are exposed and a closed position in which at least some components of the probe are received inside the body. The body is movable relative to the chassis between a high rolling position in which the body is elevated relative to the wheels and a low receiving position in which the body is lowered relative to the wheels to increase the receptiveness of the probe components so as to limit heat losses.

The dynamic arc-consistent time constraint reasoning method for deep space probe mission planning disclosed by the present invention belongs to the field of aerospace technology. The implementation method of the present invention is as follows: establish a time planning problem model for deep space probes; before performing time constraint reasoning, represent the activity time variables and time constraints as variable points and edges in a simple time network (STN), then initialize the time constraint network, and dynamically add the time variables and time constraints of newly added activities in the plan to this network; judge the constraint propagation through the variable value domain update list, only calculate the new value domain for the variables within the influence range of the new constraints, and limit the constraint propagation by whether the calculated value domain is reduced or not, narrow the scope of constraint propagation, reduce the number of time constraint reasoning times, save the limited computing resources of deep space probes, improve the efficiency of complex time constraint reasoning, improve the efficiency of deep space probe mission planning, and enhance the real-time performance of mission planning for deep space probes in case of emergencies.

1. Name of the product in this design: Launch Vehicle (Chitu 6A). 2. Purpose of this design: For use in space transportation vehicles to send payloads such as artificial Earth satellites, manned spacecraft, space stations, or space probes into predetermined space orbits. 3. The key design elements of this product are the combination of shape, pattern, and color. 4. The image or photograph that best illustrates the design's key points: a 3D model. 5. The design for which protection is sought includes color.

The application discloses a binary asteroid hovering probe adaptive control method, and belongs to the technical field of deep space probe control. The bounded periodic orbit of the probe near the equilibrium point is obtained by selecting the orbit initial value, and the probe is kept near the nominal hovering orbit by the adaptive control, so that the fuel consumption of the probe during long-time hovering can be effectively reduced; the weight matrix is dynamically selected on the basis of the linear quadratic optimal control, and at the initial moment, the probe has a large position error from the nominal position, so that the controller outputs a large control force to make the probe reach the target position as soon as possible; when reaching the vicinity of the nominal hovering orbit, the controller only needs to provide a small control force to realize the orbit keeping motion of the probe, and further reduce the fuel consumption of the probe. The position error between the nominal hovering orbit near the equilibrium point and the state of the probe is fuzzified, and the adaptive ability of the probe is improved through the weight selection.

This invention discloses a robust tracking control method for flexible attachment trajectories on small celestial bodies, belonging to the field of deep space probe control technology. The implementation method is as follows: While retaining the flexibility characteristics, the structure of the flexible probe is simplified. A "spring-damping-torsion spring" model of the flexible probe is established in both the flexible node reference coordinate system and the node fixed coordinate system, enabling it to possess both flexibility and, as far as possible, effectively estimate and control the state of its attachment process. A nominal trajectory error tracking control model is established based on the dynamic model. A performance index function is designed to transform the flexible probe attachment trajectory tracking control problem into an optimal problem. An HJB equation is established based on the performance index function, and the expression for the optimal control force is calculated. Through strategy iteration, the control law and performance index function are formed into a closed-loop equation. A single-evaluation neural network is used to approximate the optimal control law and optimal performance index function, achieving robust tracking of the nominal trajectory while reducing probe fuel consumption.

The deep space probe optical navigation position and posture decoupling estimation method disclosed by the application belongs to the field of deep space exploration. The method of the application is as follows: a navigation camera is used to image a target celestial body surface, and a pixel, a line coordinate and an image point observation error covariance matrix of a navigation landmark are extracted; according to the covariance propagation law of a nonlinear function, the image point observation error covariance is propagated to a line of sight (LOS) angle, and the observation error variance of the LOS angle is calculated; a weight is used to represent the relative accuracy between different observation angles, and an observation error weight matrix W for optical navigation position and posture decoupling estimation is constructed A ; by performing least squares weighting on the angle observation equation to incorporate the angle observation uncertainty, the weight of the navigation landmark with different observation qualities is adjusted, the angle observation error uncertainty is weighted and optimized in the absolute position estimation process, and the probe position is obtained; and the probe position and the LOS observation quantity obtained by separate decoupling estimation are used to decouple and estimate the probe posture through a multi-vector posture determination principle.

The invention belongs to the technical field of inter-satellite measurement and communication, and particularly relates to a phased array beam angle error extraction and tracking system and method and application. According to the tracking method, based on the monopulse angle measurement basic principle, a simplified phased array beam angle error extraction mathematical model is provided for the actual engineering of phased array beam angle error extraction, an appropriate threshold value is set by combining antenna characteristics, a future wave signal is converted and synthesized into an azimuth angle error and a pitch angle error, and the tracking accuracy is improved. And finally, the wave control code is sent to the phased array antenna. According to the method, the complexity of hardware FPGA design is reduced, the beam angular width parameter of the phased-array antenna can be fully utilized, the beam updating frequency is reduced, the power of a communication link can be further stabilized, the dynamic adaptability to the attitude of the deep space probe can be realized, the accurate and stable pointing of the beam under the condition of spacecraft attitude change can be realized, and the reliability of the system is improved. The problem of continuous communication of a satellite-ground communication system during large attitude jitter and large-range orbit transfer of a deep space spacecraft is solved.

The utility model relates to the technical field of space detectors, and discloses a radiation cooling plate structure for a space detector, which comprises a radiation cooling coaming plate, a notch is arranged on the front surface of the radiation cooling coaming plate, a radiation cooling plate is arranged in the notch, an installation mechanism is arranged in the radiation cooling coaming plate, and the installation mechanism is connected with the radiation cooling coaming plate. The mounting mechanism comprises a U-shaped plate, a groove, a connecting block, a clamping groove and a clamping rod, the U-shaped plate is fixedly connected to the front face of the inner wall of the radiant cooling coaming, the connecting block is fixedly connected to the back face of the radiant cooling coaming, the groove is formed in the front face of the U-shaped plate, the connecting block is inserted into the groove, the clamping groove is formed in the connecting block, and the clamping rod is fixedly connected to the clamping groove. One end of the clamping rod penetrates through the outer wall of the U-shaped plate and extends into the clamping groove; by pressing the pressing rod, the pressing rod pulls the clamping rod through the hinge rod, so that the clamping rod is separated from the clamping groove, the limitation of the connecting block is relieved, the radiation cooling plate can be conveniently disassembled, and a detector in the internal space of the radiation cooling coaming can be conveniently overhauled.

The present invention discloses a system for automatic detection of the attitude of a large space probe payload, which belongs to the field of metrology and detection in the manufacturing industry. The present invention includes a ground measurement network component, an omnidirectional mobile space positioning platform, a precise positioning platform, and a tracking and collimating measuring instrument. The present invention uses an adaptive collimating target to construct a ground measurement network with the support of a tracking and collimating measuring instrument, uses one of the dual tracking and collimating measuring instruments to cooperate with the adaptive collimating target in the ground measurement network to realize real-time measurement of the spatial position and attitude of the tracking and collimating measuring instrument, uses one of the dual tracking and collimating measuring instruments to realize alignment measurement with the cubic mirror on the satellite payload, uses a high-altitude carrying platform to realize initial positioning and stable support of the tracking and collimating measuring instrument according to the measured spatial position and attitude data, uses a tracking and collimating measuring instrument support frame to realize precise positioning of the dual tracking collimators according to the measured spatial position and attitude data, and realizes automatic measurement of the attitude of a large space probe payload.