The detecting CCD of a space astronomical telescope needs to be cooled to -75℃ to suppress the dark current for faint target detecting in the universe, and coplanarly spliced with two fine guidance sensor(FGS) which needs to be cooled to -40°C for the stability as long time observation. Two one stage thermos-electric cooler(TEC) was connected to actively cool the detector to ensure the working temperature and the temperature control accuracy, the Structural of the actively cooling detector assembly and the focal plane component were presented and the power dissipation of the TEC was calculated. In order to ensure the coplanarity of the focal plane component on the working temperature, the finite element method was used to analyze the thermal distribution on the detector surface and the thermal deformation of the supporting structure of the FGS with different materials. The analysis results showed that the lowest cooling temperature of the detecting CCD is -75°C, the temperature control accuracy was better than 1°C, and the coplanar error of the detection CCD and the fine guidance sensors did not exceed 20μm. The thermal equilibrium test showed that the lowest cooling temperature was -74.9°C~-75.1°C for the detecting CCD, The temperature control accuracy was 0.1°C. The thermal optical test showed that the defocus of the FGS was 4μm after focusing, which verified the thermal and structural design performance of the focal plane component.
A Φ1400mm silicon carbide (SiC) mirror assembly was designed according to the requirement of the mass and the optical surface distortion. The parameters of the light-weighted open-back primary mirror were optimized by finite element analysis. Six flexure bipods were designed to support the mirror edge in 12 points evenly. 12 floating anti-gravity supports were used to Minimize the optical surface distortion caused by gravity effect to obtain the real optical surface during polishing. The mirror was precisely assembled with the bipods supports and the Carbon fiber reinforcement plastic (CFRP) chamber. The optical test with interferometer showed that the surface distortion was less than 0.03λ (λ=632nm) RMS with ±5°C temperature variation and 1g gravity condition, and the mass was 145kg, which coincided with the FEA results.
The Space-based multi-band astronomical Variable Objects Monitor (SVOM) project is a dedicated satellite developed at the cooperation of China and France, aim to make prompt multi-band observations of Gamma-Ray Bursts (GRBs), the afterglows and other high-energy transient astronomical events. The Visible Telescope (VT) is one of the four payloads onboard the SVOM. VT is designed to observe the afterglows of GRBs both in the visible and near infrared bands simultaneously. The telescope can reach a limiting magnitude of +22.5Mv and provide the redshift indicators for high-Z (z<4) GRBs. VT is also designed to measure the Relative Performance Errors (RPEs) for the satellite attitude and orbit control system (AOCS), aiming to improve the pointing stability of the platform during observation. VT adopts a Ritchey-Chrétien (RC) catadioptric optical configuration with a 440mm aperture and uses the dichroic prism before the focal plane to split the incident light into blue (visible) and red (near infrared) band. Two Fine Guidance Sensor (FGS) CCDs are mounted beside the main CCD on the blue band focal plane of VT and provide sub-arcsecond pixel resolution. Fiber reinforced plastic (CFRP) composites is selected as the material of VT’s main structure to ensure enough stiffness and strength during launch. The electrical video processing circuit is carefully designed to make the readout noise below 6e-/pix (rms) in 100s exposure time. Active and passive thermal control are used together to ensure the optical performance and thermoelectric cooler (TEC) is adopted to control the main CCDs working temperature below -65°C to reduce the noise. This paper provides a comprehensive overview of the scientific requirements and the key instrument design aspects of optics, main structure, electrics, thermal control, performance test and validation results of VT.
The image fusion of optical images and synthetic aperture radar (SAR) images are of great significance. By using the complementary advantages of both, the target detection and recognition can be relatively simple and the accuracy will be relatively improved. The image information reflected by the optical image and the SAR image is very different, and the image fusion can combine the two information to give greater advantages. Aiming at the limitations of single sensor in terms of spectrum and spatial resolution, the multi-source sensor fusion technology can maximize the information description of the target scene. The fusion experiment and evaluation of optical images and SAR images are carried out by combining àtrous wavelet transform and IHS transform, and compared with the traditional HIS transform and wavelet transform fusion methods. The results show that the fusion of àtrous wavelet and HIS transform is the best, and the advantages of two single fusion methods are absorbed. It not only improves the spatial detail expression of the original image, but also preserves the spectral information of the original image, providing more accurate data for remote sensing applications.
The presence of speckle noise seriously affects the application of synthetic aperture radar (SAR) images in image fusion, so it is especially important to suppress speckle noise. According to the formation mechanism of speckle noise, this paper proposes a SAR image speckle noise removal algorithm based on improved anisotropic diffusion. The algorithm improves the diffusion coefficient c(x) based on the P-M equation diffusion filter algorithm, and adds the iterative termination condition, and obtains the filtering algorithm suitable for SAR images. This method can not only solve the problem that there are many isolated noise points in the traditional P-M model filtering, but also has a good effect on image edge preservation. The simulation results show that the improved P-M model can eliminate noise well and maintain the edge information of the image well.
Computer-generated hologram (CGH) is an effective way to compensate wavefront aberration in null test of aspheric surfaces and freeform surfaces. Our strategies of CGH design for 820mm diameter tertiary mirror surface figure testing are presented, and an experiment demonstrating the compensation test results of CGH is reported. We design a CGH including two sections on the same substrate in order to align the CGH to the incident wavefront: main section for compensating wavefront in null test, alignment section for adjusting the relative position between CGH and interferometer. Because there is no center hole in the mirror, in order to isolate different orders of diffraction, we used tilt carrier to make different orders of diffraction come to focus at different position perpendicular the axis to avoid ghost reflections.
The manufacturing and testing of a surface modified silicon carbide mirror with a bowl-shaped structure was introduced. The entire process flow includes pre-modification silicon carbide substrate processing, silicon carbide substrate surface modification, and silicon modified layer processing. Firstly, before the modification, the conventional processing method of silicon carbide was used, and the effect of the support form on the figure was eliminated by multiple direction rotation testing.At the same time, the self-aligned compensation cross-test was completed and the accuracy of the aspherical surface coefficient was verified. In addition, the polishing process of the silicon modified layer material was studied, and the optimum process parameters suitable for polishing the silicon modified layer material were found out. Based on the above experiments, the modified optical processing adopts a combination of two kinds of polishing technology: flexible chemical mechanical polishing (FCMP)and ion beam figuring (IBF).The surface roughness and surface finish of silicon modified layer are improved by flexible chemical mechanical polishing technology. The high figure accuracy of silicon modified layer is achieved finally by ion beam figuring technology. Finally, the final result of the mirror after IBF is:the RMS values of the figure and roughness in the Φ450 mm aperture is 0.01λ (λ=632.8 nm) and 0.52 nm. The mirror's processing results fully meet the design specifications.
A Φ450mm primary mirror subsystem of a space-based astronomy telescope was designed with mass, optical surface distortion and reflectivity requirement. The open-back primary mirror was made of pressure-less sintering silicon carbide, light-weighted at a ratio of approximately 70%. Three side supporting invar flexure bipods were designed to minimize the assembling stress and the thermal stress. The high reflection was obtained from the optical surface cementite. The mirror weighted 7kg and the reflectivity was 97% after optical polishing. The mirror subsystem was precisely assembled under the strict technical condition. The optical test with interferometer showed that the optical surface distortion is less than 1/40λ rms, which met the critical optical requirements for the primary mirror of the space-based astronomy telescope.
A Φ600mm SiC primary mirror subsystem of a space-borne Ritchey-Chretien telescope was designed. The open-back primary mirror was made of pressure-less sintering silicon carbide (SiC), light-weighted at a ratio of approximately 70%. Minimizing the optical surface astigmatism was critical for the mirror, the astigmatism is caused mainly by gravity effects, temperature variation and the bonding process. Three invar flexure bipods were fixed on the baseplate of the telescope at first, and the posture of the primary mirror was adjusted precisely for 0.2mm gap to the bipods. 3M 2216 B/A grey adhesive was then injected into the gap between the mirror and bipod flexure, the curing process was last 72 hours in the room temperature. So the mirror was affected only by curing stress of the adhesive during the assembly process. Structural strength and dynamic stiffness of the mirror subsystem in the thermal- structural coupling state were analyzed with finite element method. Analyzed results show that the optical surface distortion is less than 1/50λ at 632.8nm RMS with three points support and less than 1/200λ RMS with 2°C temperature variation because of the flexure support and compatible support and mirror material, The optical performance test with interferometer show that the optical surface distortion caused by the curing stress of the adhesive is less than 1/50λRMS, the overall optical surface of the primary mirror is less than 1/30λ rms, which met the critical requirements for the primary mirror of the telescope.
One space-based astronomy telescope will observe astronomy objects whose brightness should be lower than 23th magnitude. To ensure the telescope performance, very low system noise requirements need extreme low CCD operating temperature (lower than -65°C). Because the satellite will be launched in a low earth orbit, inevitable space external heat fluxes will result in a high radiator sink temperature (higher than -65°C). Only passive measures can’t meet the focal plane cooling specification and active cooling technologies must be utilized. Based on detailed analysis on thermal environment of the telescope and thermal characteristics of focal plane assembly (FPA), active cooling system which is based on thermo-electric cooler (TEC) and heat rejection system (HRS) which is based on flexible heat pipe and radiator have been designed. Power consumption of TECs is dependent on the heat pumped requirements and its hot side temperature. Heat rejection capability of HRS is mainly dependent on the radiator size and temperature. To compromise TEC power consumption and the radiator size requirement, thermal design of FPA must be optimized. Parasitic heat loads on the detector is minimized to reduce the heat pumped demands of TECs and its power consumption. Thermal resistance of heat rejection system is minimized to reject the heat dissipation of TECs from the hot side to the radiator efficiently. The size and surface coating of radiator are optimized to compromise heat reject ion requirements and system constraints. Based on above work, transient thermal analysis of FPA is performed. FPA prototype model has been developed and thermal vacuum/balance test has been accomplished. From the test, temperature of key parts and working parameters of TECs in extreme cases have been acquired. Test results show that CCD can be controlled below -65°C and all parts worked well during the test. All of these verified the thermal design of FPA and some lessons will be presented in this paper.
Mechanical stability is a significant segment for an on-axis space telescope to assure its assembly accuracy as well as the image quality in the rigorous space environment, supporting structure between the primary mirror and the secondary mirror as a main structure of the on-axis space telescope must be designed reasonably to meet the mission requirements of the space telescope. Meanwhile, in view of the limitation of the satellite launching cost, it is necessary to reduce the weight and power compensation during the supporting structure design based on the satisfaction of telescope performance. Two types of supporting structure for a space telescope are designed, one is three-tripod structure which has three tripods located on the optical bench to support the secondary mirror assemblies and keep the distance between the primary mirror and the secondary mirror, the other is barrel supporting structure which includes a tube and a secondary mirror support with four spider struts. To compare the mechanical performance and launching cost of the two kinds of supporting structure, both structural and thermal analysis model are established. The analysis results indicates that the three-tripod support is lighter, has better mechanical performance and needs less power compensation than the barrel support.
An attitude-varied space camera changes attitude continually when it is working, its attitude changes with large angle in short time leads to the significant change of heat flux; Moreover, the complicated inner heat sources, other payloads and the satellite platform will also bring thermal coupling effects to the space camera. According to a space camera which is located on a two dimensional rotating platform, detailed thermal design is accomplished by means of thermal isolation, thermal transmission and temperature compensation, etc. Then the ultimate simulation cases of both high temperature and low temperature are chosen considering the obscuration of the satellite platform and other payloads, and also the heat flux analysis of light entrance and radiator surface of the camera. NEVEDA and SindaG are used to establish the simulation model of the camera and the analysis is carried out. The results indicate that, under both passive and active thermal control, the temperature of optical components is 20±1°C,both their radial and axial temperature gradient are less than 0.3°C, while the temperature of the main structural components is 20±2°C, and the temperature fluctuation of the focal plane assemblies is 3.0-9.5°C The simulation shows that the thermal control system can meet the need of the mission, and the thermal design is efficient and reasonable.
The limb UV radiation detection provides the information of atmosphere ultraviolet spectrum, so as to acquire the high resolution vertical distribution information of atmospheric trace gases and aerosol. Off-axis Three-mirror Anastigmat
(TMA) system is adopted in limb UV radiation detection to increase horizontal space coverage. In this paper, opto-
mechanical design of the system is introduced, and camera alignment is completed by computer aiding, then optical,
mechanical and electrical combination as well as the optical performance test are carried out with the UV Image Intensifier. The camera’s wavefront error of each field is close to design value after alignment, the resolution reaches
140lp/mm in visible light band, and 20lp/mm in UV band, which reaches the design limit of the UV Image Intensifier, the optical system could well meet the operational requirement.
A space telescope containing two CCD cameras is being built for scientific observation. The CCD detectors need to
operate at a temperature below -65°C in order to avoid unacceptable dark current. This cooling is achieved through
detailed thermal design which minimizes the parasitic load to 2K×4K array with 13.5 micron pixels and cools this
detector with a combination of thermo electric cooler(TEC).
This paper will describe detailed thermal design necessary to maintain the CCD at its cold operating temperature while
providing the means to reject the heat generated by the TECs. It will focus on optimized techniques developed to manage
parasitic loads including material selection, surface finishes and thermal insulation. The paper will also address analytical
techniques developed to characterize TEC performance. Finally, analysis results have been shown the temperature of key
parts.
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