Radiation sensor for deep space environments

Abstract

To provide a radiation sensor for using in a deep space environment.SOLUTION: The present invention provides a radiation sensor for the deep space environment including: a circuit substrate; a payload control module disposed on the circuit substrate; a radiation sensitive field-effect transistor readout module that is disposed on the circuit substrate, and electrically connected to the payload control module; and a flash memory that is disposed on the circuit substrate, and electrically connected to both the payload control module and the radiation sensitive field-effect transistor readout module. In the flash memory, detection software is included that, upon detecting a single event upset in at least one bit of stored sensor data, immediately identifies and records a position of an affected bit. Then, the detection software resets affected data to record the number of bit errors.SELECTED DRAWING: None

Description

[Technical Field]

[0001] The present invention relates to a radiation sensor for a deep space environment, and more particularly to a radiation sensor for a deep space environment that can withstand high-intensity radiation, high-energy particles, extreme temperatures, heavy mechanical loads, and severe vibrations. [Background technology]

[0002] Deep space exploration refers to space exploration missions that leave low Earth orbit (below 2000 kilometers) and enter high Earth orbit, interplanetary space, or near-Earth space, and typically involve exploration of the Moon or other celestial bodies. The deep space environment is significantly different from the near-Earth orbit environment, and due to the longer duration of exploration in deep space, deep space sensors may be exposed to space conditions in deep space orbit that are worse than those in near-Earth orbit.

[0003] The deep space environment is filled with influences from high-energy radiation fields, solar energetic particles (SEPs), photons, galactic cosmic rays (GCRs), dense, high-energy plasma due to coronal mass ejections, and induced radiation fields on the surfaces of celestial bodies, all of which can degrade or disable the performance of functional materials and components in deep space sensors and can cause physiological reactions in astronauts, such as cataracts, neurological disorders, and weakened immune systems.

[0004] In the deep space environment, ionizing radiation poses a significant risk to sensors and astronaut safety during space travel. The deep space environment extends beyond Earth's magnetosphere and lacks the protection of Earth's magnetic field and atmosphere, resulting in a high flux of ionizing electromagnetic and particle radiation. Based on the sensor's flight in the deep space environment, the ionizing radiation environment can be divided into three domains: (1) the cosmic radiation environment during the journey from Earth to other planets, which is primarily due to solar energetic particles and galactic cosmic rays; (2) the radiation environment during the sensor's descent to a deep space body, which is primarily due to solar energetic particles and galactic cosmic rays captured by the body's magnetic field; and (3) the radiation environment on the surface of a deep space body where the sensor lands, which is primarily composed of secondary radiation generated after the body absorbs cosmic radiation, primarily consisting of high-energy, heavy particles.

[0005] Astronauts' bodies are exposed to radiation from the deep space environment, and the longer they remain in the deep space radiation environment, the greater the effects. Absorption of high-energy particles, photons, and cosmic rays damages astronauts' cells, causing DNA mutations and increasing their risk of cancer, as exposure disrupts the molecular structures within most living organisms. Radiation penetrates living tissues, causing short- and long-term damage to bone marrow stem cells and leading to chromosomal abnormalities in lymphocytes, the core of the immune system. Damage can weaken immunity to viruses previously suppressed by the body. Furthermore, lymphocyte T cells are less likely to reproduce normally in space, and even if they do, they have difficulty resisting infection. This reduces astronauts' resistance to disease, and the confined space of a spacecraft increases the risk of disease transmission among crew members.

[0006] In the deep space environment, sensors encounter high-energy charged particle radiation from solar energetic particles and galactic cosmic rays, as well as high-energy solar electromagnetic radiation. When sensors operate near planets or their satellites, they may be affected by low-energy charged particle radiation from induced radiation belts and neutron radiation from the surface of celestial bodies. The harsh radiation environment of deep space has more severe degrading effects on sensitive sensor materials and components than those experienced in low Earth orbit. Long-term exposure to the deep space radiation environment significantly degrades the performance of sensitive materials within sensors, including thermal control materials, solar cells, optical materials, insulating materials, and sealing materials (including degradation of optical, electrical, and mechanical properties, as well as insulation and sealing properties). Electronic components are also susceptible to single-event effects and total ionizing dose effects.

[0007] A single event effect (SEE) is a series of abnormal effects caused by the passage of a high-energy charged particle through a microelectronic component, resulting in abnormal changes or damage to the component's logic state or function due to charge collection at the electrodes of sensitive components. SEEs can alter a computer's logic bit(s) and cause permanent, destructive damage. The occurrence of an SEE is a matter of probability with each high-energy particle strike. Long-term radiation exposure, known as total ionizing dose (TID), can also gradually change the performance of electronic products. If a circuit receives a certain level of total ionizing dose, it will cease to function properly, eventually leading to permanent failure.

[0008] A common class of SEE is the single-event upset (SEU), which is a single high-energy particle applied to a semiconductor component, causing an abnormal change in the component's logic state. SEUs are the most common and typical of the various SEEs caused by space radiation, occurring primarily in data storage or instruction-related components. Component errors caused by SEUs are considered "soft errors" and can be restored to their normal state by a system reset, power cycle, or rewrite. However, SEUs can also cause data corruption, abnormal software state transitions, or software hangs or crashes.

[0009] Another common class of SEEs is single-event latch-up (SEL), a phenomenon in which a single high-energy particle penetrates a semiconductor structure, mutating parasitic structures into a low-impedance state and creating overcurrents that can permanently lose their functionality if not immediately powered back on. This primarily affects CMOS components. For example, in a silicon-controlled rectifier, originally a PNPN four-layer structure, a single charged particle can induce a transient current, causing the silicon structure to conduct and decompose into stacked PNP and NPN transistors. Single-event latch-up can cause irreparable damage to electronic components. Unlike SEUs, this is considered a "hard error" that causes physical hardware damage, making repairs in space even more difficult, posing a significant challenge to space radiation detection.

[0010] Therefore, both astronauts and electronic components are subject to high-energy radiation from the deep space environment. As a result, providing radiation sensors that can withstand and recover from SELs, SEUs, and TIDs for use in deep space environments is a challenging problem that engineers in this field are trying to solve. Summary of the Invention [Problem to be solved by the invention]

[0011] It is an object of the present invention to provide a radiation sensor for use in deep space environments that can operate under conditions of high intensity radiation, high energy particles, extreme temperatures, severe vibrations, and mechanical loads, that is capable of measuring ionizing radiation dose, dose rate, and number of single event upsets in deep space, and that is immune to SEUs and single event latch-ups. [Means for solving the problem]

[0012] To achieve the above object, the present invention provides a radiation sensor for deep space environments, including a circuit board, a payload control module disposed on the circuit board, a radiation-sensitive field-effect transistor readout module disposed on the circuit board and electrically connected to the payload control module, and a flash memory integrated circuit (IC) disposed on the circuit board and electrically connected to both the payload control module and the radiation-sensitive field-effect transistor readout module. The flash memory contains detection software. The flash memory stores sensor data under an ionizing radiation environment of 0 to 100,000 rad and detects at least one bit flip upon an SEU. The detection software instantly detects and records the location of the bit in the flash memory where the SEU occurred, resets the data where the SEU occurred, and records the number of bit errors. By sweeping and resetting bits affected by the SEU, the detection software is configured to sense the effects of cosmic ray heavy ions and solar energetic particles on the flash memory and associated computer memory. This may be useful for future spacecraft avionics designs.

[0013] In one embodiment of the present invention, the electronic device further includes a chassis arranged around the outside of the circuit board and having a through hole on one side and a first cutout adjacent to the through hole, a front panel member arranged on the chassis and having a second cutout, and at least one fixing member arranged between the front panel member and the circuit board and configured to fix the circuit board and separate the circuit board and the front panel member by a certain distance.

[0014] In one embodiment of the present invention, the device further includes a plurality of insulating collars disposed on the front panel member.

[0015] In one embodiment of the present invention, the device further includes an electrical interface disposed on the circuit board for electrically connecting the payload control module and the radiation sensitive field effect transistor readout module, and a data interface disposed on the circuit board for electrically connecting the payload control module and the radiation sensitive field effect transistor readout module.

[0016] In one embodiment of the present invention, the electrical interface further includes a transformer electrically connected to the payload control module, a first electronic fuse electrically connected to the radiation sensitive field effect transistor readout module, and a second electronic fuse electrically connected to the transformer, wherein if a single event latch-up occurs in the radiation sensitive field effect transistor readout module in the ionizing radiation environment, it is cleared by cycling power using either the first or second electronic fuse.

[0017] In one embodiment of the present invention, the first electronic fuse is electrically connected to a first input power source, and the second electronic fuse is electrically connected to a second input power source.

[0018] In one embodiment of the present invention, the circuit board is disposed in an aircraft, a rocket, or a satellite, respectively.

[0019] In one embodiment of the present invention, the front plate members are disposed on the surface of an aircraft, a rocket, or a satellite, respectively.

[0020] One embodiment of the present invention includes a multi-layer insulating member disposed outside the chassis, surrounding the chassis, and configured for temperature regulation.

[0021] In one embodiment of the present invention, the top plate of the chassis is coated with a white UV-resistant coating and functions as a heat-dissipating front panel member.

[0022] To this end, a radiation sensor for use in deep space environments is provided, which solves problems that space science experts want to solve. [Brief explanation of the drawings]

[0023] [Figure 1] FIG. 1 is a block diagram of one embodiment of the present invention.

[0024] [Figure 2] 1 is a system block diagram of an embodiment of the present invention.

[0025] [Figure 3] FIG. 1 is a block diagram of one embodiment of the present invention.

[0026] [Figure 4] 1 is a top schematic view of the present invention.

[0027] [Figure 5] FIG. 2 is a structural diagram of a front plate member of the present invention.

[0028] [Figure 6] 1 is a side schematic view of the present invention.

[0029] [Figure 7] FIG. 10 is a comparative diagram of random vibration power spectral densities of the present invention.

[0030] [Figure 8A] FIG. 1 is a schematic diagram of the X-axis and Y-axis sine wave test of the present invention.

[0031] [Figure 8B] FIG. 1 is a schematic diagram of a Z-axis sine wave test of the present invention.

[0032] [Figure 9] FIG. 1 is a schematic diagram of a thermal test of the present invention.

[0033] [Figure 10] FIG. 1 is a schematic diagram of radiation test results of the present invention.

[0034] [Figure 11] 1 is a schematic diagram showing the results of a proton beam irradiation test of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0035] To provide a further understanding and appreciation of the features and advantages achieved by the present invention, preferred embodiments are presented below along with detailed descriptions.

[0036] Because the deep space environment is filled with high-energy mixed cosmic radiation fields, detecting radiation dose is an important part of deep space exploration or detection. However, high-energy radiation can cause errors in electronic components, preventing radiation sensors from accurately measuring radiation doses in deep space. Inaccurate radiation doses can easily lead to incorrect decisions by astronauts and central control computers, which can pose a danger to astronauts and spacecraft.

[0037] The radiation sensor of the present invention, which is applied to a deep space environment, can not only withstand high-intensity radiation, high-energy particles, extreme temperatures, and severe vibrations, but can also detect and repair SEUs in electronic components while measuring the radiation dose, thereby preventing accidents caused by erroneous radiation dose measurements and control sequence disruption due to SEUs in a deep space environment.

[0038] Various embodiments of the present invention will now be described in detail with reference to the figures. However, the concept of the present invention may be embodied in various forms and should not be construed as being limited to the exemplary embodiments set forth herein.

[0039] First, the circuit portion of the present invention will be described. Referring to FIG. 1, FIG. 1 is a block diagram of one embodiment of the present invention. Next, the radiation sensor 1 for deep space environments will be described in detail. The radiation sensor 1 for deep space environments includes a circuit board 10, a payload control module 20 disposed on the circuit board 10, a radiation-sensitive field-effect transistor readout module 30 disposed on the circuit board 10 and electrically connected to the payload control module 20, and a flash memory 40 disposed on the circuit board 10 and electrically connected to both the payload control module 20 and the radiation-sensitive field-effect transistor readout module 30. The flash memory 40 includes detection software 42, which stores detection data under a high-ionizing radiation environment ranging from 0 to 100,000 rad. When at least one bit of the detection data experiences an SEU, the detection software 42 instantly detects and records the location of the bit in the flash memory 40 where the SEU occurred. The detection software 42 resets the data where the SEU occurred and records the number of bit errors.

[0040] Continuing from the above, the present invention senses radiation in a deep space environment via a radiation-sensitive field-effect transistor readout module 30. The radiation-sensitive field-effect transistor readout module 30 includes a radiation-sensitive field-effect transistor 32, which is a P-channel MOSFET optimized for radiation sensitivity. The radiation-sensitive field-effect transistor 32 is specially designed to be sensitive to high-energy (ionizing) radiation and exhibits a threshold voltage that varies depending on the amount of ionizing radiation absorbed. The present invention involves recording the voltage of this type of microelectronic chip at specified time intervals and calculating the amount of ionizing radiation from the recorded threshold voltage. The calculation method can be improved by comparing the recorded voltage with calibration data obtained under simulated laboratory conditions, thereby estimating the amount of ionizing radiation, but is not limited to this method.

[0041] Following on from the above, the present invention utilizes detection software 42 in flash memory 40 to detect SEUs while also using flash memory 40 for data storage. Flash memory 40 not only has excellent dynamic shock resistance, preventing data loss due to violent shaking, but is also very robust when used as a memory card, able to withstand high pressures and extreme temperatures. Furthermore, as non-volatile solid-state storage, it consumes no power while saving files, providing many advantages for deep space exploration.

[0042] However, even without the effects of ionizing radiation, inaccuracies can occur in flash memory 40 due to data retention errors or interference with read / write operations. Therefore, it is necessary to perform periodic error detection and correction on flash memory 40. One correction and two error detections per word are believed to be sufficient for flash memory 40.

[0043] Following the above, in addition to its data storage function, the flash memory 40, when exposed to high ionizing radiation in outer space, encounters SEUs caused by cosmic ray heavy ions or solar energetic particles. As a result, the detection software 42 in the flash memory 40 quickly detects and records the location of bits affected by the SEU. The detection software 42 resets the data affected by the SEU and records the number of bit errors.

[0044] Following the above, the present invention controls the payload via the payload control module 20. When the payload control module 20 is activated, the payload enters a science mode and periodically generates and records a number of management data packets. The packets include health status data of the payload and the last recorded science data. The health status data corresponds to whether there is an error in the software of the payload control module 20, and the science data corresponds to the voltage change detected by the radiation sensor field effect transistor 32 and the number of SEUs detected by the detection software 42 in the flash memory 40.

[0045] Please refer to Fig. 2, which is a system block diagram of one embodiment of the present invention. As shown in the figure, the present invention is applied to a radiation sensor 1 for a deep space environment. The radiation sensor 1 for a deep space environment further includes an electrical interface 50 disposed on a circuit board 10 and electrically connected to a payload control module 20 and a radiation-sensitive field-effect transistor readout module 30, and a data interface 60 disposed on the circuit board 10 and electrically connected to the payload control module 20 and the radiation-sensitive field-effect transistor readout module 30.

[0046] Continuing from the above, the electrical interface 50 further includes a transformer 52 electrically connected to the payload control module 20, a first electronic fuse 54 electrically connected to the radiation sensitive field effect transistor readout module 30, and a second electronic fuse 56 electrically connected to the transformer 52. The first electronic fuse 54 and the second electronic fuse 56 are configured to restore the SEL.

[0047] Following on from the above, the present invention provides a mechanism for autonomously recovering the SEL via the first electronic fuse 54 and the second electronic fuse 56. The SEL causes a high current to flow through electronic components, causing a fault that can only be cleared by restarting the power supply. Therefore, when the first electronic fuse 54 or the second electronic fuse 56 detects an overcurrent, it temporarily cuts off the power supply to facilitate restarting the power supply to recover from the SEL.

[0048] Continuing from the above, when applied to a spacecraft, rocket, or satellite, the radiation sensor 1 of the present invention is connected to an external payload interface card 70 of the spacecraft, rocket, or satellite. The data interface 60 of the present invention is configured to convert the data protocol of the payload control module 20 into the data protocol of the external payload interface card 70, and can transfer data between the payload control module 20 and the external payload interface card 70.

[0049] Following the above, when configured in a deep space environment, the radiation sensor 1 of the present invention can be placed in a spacecraft, rocket, or satellite and interconnected with their own systems to achieve the effect of radiation detection.

[0050] In this embodiment, the data interface 60 comprises a transceiver integrated circuit (not shown) and supporting circuitry (not shown), but is not limited to this configuration.

[0051] Continuing from the above, reference is now made to FIG. 3, which is a schematic diagram of an external payload interface card of the present invention. As shown in the figure, the radiation sensor for deep space environments 1 of the present invention is connected to an external payload interface card 70 of a spacecraft, rocket, or satellite. The external payload interface card 70 further includes a first input power source 72 electrically connected to the first electronic fuse 54, a second input power source 74 electrically connected to the second electronic fuse 56, and a ground terminal 78 electrically connected to the data interface 60, the payload control module 20, the radiation sensitive field effect transistor readout module 30, and the flash memory 40.

[0052] Following from the above, the present invention provides a mechanism for autonomously recovering from the SEL via the first electronic fuse 54 and the second electronic fuse 56. The SEL causes a high current to flow through the electronic components, causing a failure that can only be cleared by a power cycle. Thus, when the first electronic fuse 54 or the second electronic fuse 56 detects an overcurrent, it temporarily disconnects the first input power source 72 or the second input power source 74, facilitating a power cycle of the first input power source 72 or the second input power source 74 to recover from the SEL.

[0053] In this embodiment, the first input power supply 72 is 12 V and is used as the power supply for the radiation sensitive field effect transistor readout module 30. Since the input voltage required for the radiation sensitive field effect transistor readout module 30 is 12 V, the transformer 52 for adjusting the voltage is not required. On the other hand, the second input power supply 74 is 5 V, and the voltage of the second input power supply 74 is adjusted to 3.3 V via the transformer 52 and used as the power supply for the payload control module 20. Here, 3.3 V is the input voltage required for most of the electronic components in this embodiment, except for the radiation sensitive field effect transistor readout module 30. In this example, the input voltage of the components of the present invention is not particularly limited.

[0054] In this embodiment, when the first electronic fuse 54 or the second electronic fuse 56 detects an overcurrent, it disconnects the 12V first input power supply 72 or the 5V second input power supply 74, preventing the radiation sensor 1 for deep space environments from being damaged or destroyed by a high overcurrent due to the SEL, thereby protecting the radiation sensor 1 for deep space environments.

[0055] In this embodiment, flash memory 40 is electrically connected to a data interface 60, which can read and store data stored in flash memory 40 for future retrieval.

[0056] In this embodiment, the payload control module 20 further includes a temperature sensing circuit 22 that measures both the temperature of the payload control module 20 itself and the temperature of the radiation sensitive field effect transistor readout module 30 to prevent radiation-unrelated voltage fluctuations in the radiation sensitive field effect transistor readout module 30.

[0057] Next, the structural components of the present invention will be described. Reference will be made to FIGS. 4, 5, and 6. FIG. 4 is a schematic top view of the present invention, FIG. 5 is a schematic view of the substrate structure, and FIG. 6 is a schematic side view of the present invention. As shown in the figures, the present invention is applied to a radiation sensor 1 for deep space environments. The radiation sensor 1 for deep space environments further includes a chassis 5 having a through-hole (not shown) on one side and a first recess 9 adjacent to the through-hole. A circuit board 10 is disposed within the chassis 5. A payload control module 20 is disposed on the circuit board 10. A radiation-sensitive field-effect transistor readout module 30 is disposed on the circuit board 10 and electrically connected to the payload control module 20. A flash memory 40 is disposed on the circuit board 10 and electrically connected to both the payload control module 20 and the radiation-sensitive field-effect transistor readout module 30. A front plate member 80 is disposed on the chassis 5 and has a second hollow portion 84. At least one fixing member 90 is disposed between the front plate member 80 and the circuit board 10. At least one fixing member 90 is configured to fix the circuit board 10 and separate the circuit board 10 and the front plate member 80 by a first distance D1. The radiation-sensitive field-effect transistor readout module 30 senses radiation intensity in a high-ionizing radiation environment and increases its threshold voltage. The flash memory 40 detects and records the threshold voltage of the radiation-sensitive field-effect transistor readout module 30 and the number of SEUs that have occurred in the flash memory 40. The flash memory 40 resets the data indicating the location where the SEU occurred. The present invention protects the circuit board 10, payload control module 20, radiation-sensitive field-effect transistor readout module 30, and flash memory 40 by placing the circuit board 10 in a chassis 5. At the same time, the chassis 5 provides thermal insulation and vibration protection for the circuit board 10, payload control module 20, radiation-sensitive field-effect transistor readout module 30, and flash memory 40.

[0058] In this embodiment, since the radiation sensor 1 for deep space environments of the present invention is directly exposed to outer space, it is necessary to maintain the temperatures of all electronic components within a load range to prevent malfunction of the electronic components due to temperature changes. Furthermore, it is essential to ensure that the total mass of the radiation sensor 1 for deep space environments does not exceed 400 grams and that the total power consumption is less than 900 milliwatts. Regarding temperature, since the spacecraft has a side facing the sun and a side away from the sun, the side facing the sun continues to heat up and the side away from the sun continues to cool, resulting in more extreme temperature differences.

[0059] 5, in this example, a front plate member 80 is disposed on the chassis 5, and a plurality of insulating collars 82 are disposed on the front plate member 80, as shown. The placement of the insulating collars 82 further reduces the effects of cold temperatures being conducted between the internal electronic components and the spacecraft.

[0060] Continuing from the above, the radiation sensor 1 for deep space environments of the present invention can be placed on the exterior of a spacecraft, rocket, or satellite and interconnected with a unique system to exert its radiation sensor effect. Since the exterior of a spacecraft, rocket, or satellite is a harsh environment, the chassis 5 and at least one fixing part 90 are configured to fix and protect the circuit board 10, prevent physical damage, and reduce the effects of vibration and temperature on the radiation sensor 1 for deep space environments.

[0061] Since the internal electronic components may not function properly in an environment below 0 degrees Celsius, this embodiment further includes an external heater (not shown). When the temperature detection circuit 22 detects a temperature below 0 degrees Celsius, the external heater is activated to cope with the harsh low-temperature environment and allow the internal electronic components to function properly.

[0062] Since the radiation sensor 1 for deep space environments must be able to withstand vibrations caused by the launch and flight of a spacecraft, in this embodiment, the maximum static load coefficient required is 33 G on the Z axis and 30 G on the XY plane, where G represents the gravitational acceleration, which is 9.8 meters per second squared.

[0063] Due to spacecraft equipment regulations, in this embodiment, the weight of the radiation sensor 1 for deep space environments must be less than 400 grams. Therefore, referring to Figures 4 and 5, in this embodiment, a plurality of first hollow portions 9 and a plurality of second hollow portions 84 are designed on the chassis 5 and the front plate member 80, and are fixed with several screws to provide mechanical loads in a flight environment.

[0064] Next, refer to Figure 6. As shown in the figure, the radiation sensor for a deep space environment 1 further includes at least one fixing component 90 disposed between the front plate member 80 and the circuit board 10. The at least one fixing component 90 is configured to fix the circuit board 10 and separate the circuit board 10 and the front plate member 80 by a first distance D1 to prevent the circuit board 10 from coming into direct contact with the front plate member 80, and also functions as a heat insulator.

[0065] Continuing from the above, the present invention can also incorporate a multi-layer insulating member disposed outside the chassis 5 and surrounding the chassis 5. The multi-layer insulating member is configured to regulate temperature. In this embodiment, the multi-layer insulating member can be stacked in multiple layers to enhance the temperature regulation effect. This can prevent electronic components from reaching an operating temperature due to excessively high or low temperatures, which can cause the electronic components to malfunction or be damaged.

[0066] Following the above, the present invention can also provide a UV-resistant white coating on the surface of the chassis 5, which functions as a heat sink and a radiation front panel member of the radiation sensor 1 in a deep space environment, to prevent the electronic components from reaching an operating temperature due to excessively high or low temperatures, which can cause the electronic components to malfunction or be damaged.

[0067] Following the above, please refer to the following table (1). JPEG0007822085000001.jpg130160

[0068] Please refer to Table (1) and Figure 7, which is a comparison diagram of the random vibration power spectrum density of the present invention. The upper line 102 represents the random vibration power spectrum density of the present invention, and the lower line 104 is provided by SpaceX. The maximum displacement of the present invention is 3.76 x 10 -2 It was observed that the vibrations were within the maximum allowable range of millimeters, which proves that the radiation sensor 1 for use in deep space environments of the present invention can withstand the vibrations that occur during the launch and flight of a spacecraft.

[0069] Please refer to Table (1), FIGS. 8A, and 8B. FIG. 8A shows a sine wave test diagram for the X-axis and Y-axis of the present invention, and FIG. 8B shows a sine wave test diagram for the Z-axis. As shown in the figures, a resonance test is performed on the radiation sensor 1 for deep space environments of the present invention. The sensor is fixed in different axial directions on a vibration table using a fixture, and sinusoidal vibrations are performed as shown in FIGS. 8A and 8B. A sine wave test with an amplitude of 33 G is performed along the XY plane, and a sine wave test with a test amplitude of 30 G is performed along the Z-axis. In this embodiment, the radiation sensor 1 for deep space environments was subjected to four different amplitude tests in three axes.

[0070] Following the above, after testing, the radiation sensor 1 for deep space environments of the present invention was confirmed to be 3 dB lower than the passing specification and passed the functional test after power-on. The radiation sensor 1 for deep space environments of the present invention can withstand not only the vibrations that occur during launch or flight, but also the mechanical stresses encountered during flight, and continues to operate normally even after enduring such mechanical stresses.

[0071] Continuing from the above, reference is now made to FIG. 9, which is a schematic diagram of a thermal test of the present invention. As shown in the figure, a temperature check of the radiation sensor 1 configured for a deep space environment was performed, focusing on the temperature-sensitive radiation-sensing field-effect transistor 32. As shown, this test involved a continuous four-hour hot soak at 60°C and a cold soak at 0°C to perform thermal testing of the radiation-sensing field-effect transistor 32. Functional tests were performed before and after the soaks, including the transmission and verification of a series of commands, including telemetry requests and operational mode changes. During the hot and cold soak periods, limited functional tests were performed, transmitting and verifying abbreviated commands, and sending these management data packet request commands every 30 seconds to verify that functionality remained intact.

[0072] Test results show that during the cool-down phase, the radiation-sensitive field-effect transistor 32 responds slowly to commands and exhibits an abnormal increase in current. The abnormality occurs only when the temperature approaches or drops below the ambient temperature of the thermal vacuum chamber. To restore normal function, the power must be turned off, allowed to warm up, and then reconnected at room temperature.

[0073] Since the present invention requires radiation detection between high-intensity radiation, the radiation-sensitive field-effect transistor readout module 30 must be able to withstand 10,000 rad of radiation through a 1 mm aluminum shield. Therefore, a radiation test is performed on the radiation sensor 1 configured for a deep space environment. Referring to FIG. 10, FIG. 10 is a schematic diagram of the radiation test results of the present invention. As shown in the figure, the present invention tests two sets of radiation-sensitive field-effect transistor readout modules 30, designated as lines 114 and 116, with line 112 as the standard radiation value. When the radiation dose in the radiation test is 10,168 rad, it is observed that the errors of lines 114 and 116 under radiation exposure are very small. Therefore, it is clear that the present invention can accurately detect radiation with minimal error even under a radiation intensity of 10,000 rad.

[0074] Next, it is necessary to test whether the present invention can withstand and resolve SEE. To this end, a proton beam generated by a cyclotron is irradiated onto a radiation sensor 1 configured in a deep space environment, as shown in FIG. 11. FIG. 11 is a schematic diagram of the proton beam irradiation test results of the present invention. As shown in the figure, the DSRP value represents the number of bit errors caused by proton beam irradiation. The average proton energy of the proton beam generated by the cyclotron is about 221.2 MeV, and the average proton flux is 1.32 × 10 7 cm 2 s -1 , the test duration was 33 minutes, and the proton beam flux was 2.37 × 10 10 cm -2 The radiation dose was 1290 rad. No abnormal non-commanded recovery or power cycle occurred during the test, indicating that no overcurrent due to single event latch-up occurred, proving that the deep space environment radiation sensor 1 can withstand proton impacts. Multiple peaks may be observed, but these are likely data errors caused by SEU due to proton impacts, and these can also be resolved by the detection software 42 in the flash memory 40.

[0075] In the above embodiment, the present invention provides a radiation sensor for use in deep space environments. Detection software in flash memory detects and repairs SEUs, and a radiation-sensitive field-effect transistor readout module records the detected radiation dose. Multiple electronic fuses are configured to resolve overcurrents caused by single-event latch-up. By utilizing a series of lightweight components to reduce weight without affecting mechanical strength, and by using multi-layer insulation for thermal insulation, the radiation sensor can operate normally in extreme temperatures, strong vibrations, and excessive radiation environments without being affected by SEUs, providing astronauts and space system engineers with reliable radiation detection data.

Claims

1. A radiation sensor for a deep space environment, comprising: A circuit board; a payload control module disposed on the circuit board; a radiation sensitive field effect transistor readout module disposed on the circuit board and electrically connected to the payload control module; a flash memory disposed on the circuit board and electrically connected to the payload control module and the radiation sensitive field effect transistor readout module; the flash memory contains detection software; the flash memory stores detected data under an ionizing radiation environment of 0 rad to 100,000 rad, and when a single event upset occurs in at least one bit of the detected data, the detection software instantly detects and records the position of the bit in the flash memory where the single event upset occurred, resets the data where the single event upset occurred, and records the number of bit errors; Radiation sensor for deep space environments.

2. a chassis disposed around the outer periphery of the circuit board, the chassis having a through hole on one side and a first recessed portion adjacent to the through hole; a front panel member disposed on the chassis and having a second recess; at least one fixing member disposed between the front plate member and the circuit board, the fixing member configured to fix the circuit board and separate the circuit board and the front plate member by a first distance; The radiation sensor for a deep space environment according to claim 1 , further comprising:

3. The radiation sensor for a deep space environment according to claim 2 , further comprising a plurality of thermal insulating collars disposed on said front plate member.

4. an electrical interface disposed on the circuit board and electrically connected to the payload control module and the radiation sensitive field effect transistor readout module; a data interface disposed on the circuit board and electrically connected to the payload control module and the radiation sensitive field effect transistor readout module; The radiation sensor for a deep space environment according to claim 1 , further comprising:

5. The electrical interface includes: a transformer electrically connected to the payload control module; a first electronic fuse electrically connected to the radiation sensitive field effect transistor readout module; a second electronic fuse electrically connected to the transformer; further comprising the radiation sensitive field effect transistor readout module is configured to sense a radiation dose due to the ionizing radiation environment, and the first or second electronic fuse is configured to recover from a single event latch-up due to the ionizing radiation environment through a power cycle; The radiation sensor for a deep space environment according to claim 4 .

6. 6. The radiation sensor for a deep space environment according to claim 5, wherein the first electronic fuse is electrically connected to a first input power source, and the second electronic fuse is electrically connected to a second input power source.

7. The radiation sensor for a deep space environment according to claim 1 , wherein the circuit board is disposed in an aircraft, a rocket, or an artificial satellite, respectively.

8. The radiation sensor for a deep space environment according to claim 2 , wherein the front plate member is disposed on the surface of an aircraft, a rocket, or an artificial satellite.

9. The radiation sensor for deep space environments according to claim 2 , further comprising a multi-layer thermal insulation member configured to cover and surround an exterior of the chassis and to regulate a temperature within the chassis.

10. 3. The radiation sensor for use in a deep space environment according to claim 2, wherein a surface of the chassis is coated with an ultraviolet resistant white coating.

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