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.
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.
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.
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
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.