尖化前缘热环境实验技术研究
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摘要
随着新型高超声速飞行器的发展,能够保持长时间、长航程飞行的高升阻比气动外形成为这类飞行器的技术关键。由于钝化前缘已无法适应新型飞行器的气动性能要求,出现了低阻力的尖化前缘外形飞行器,其在气动性能上存在较大优势,但会带来热流密度大、防热困难等问题。另外,新型高超声速飞行器要保持外形,维持高升阻比,必需采用非烧蚀的热防护技术,对热环境预测精度提出更高的要求。开展尖化前缘热环境实验技术研究,为尖化前缘的热环境理论预测和热防护提供可靠的实验数据,对于新型高超声速飞行器的研制具有重要意义。
本文通过对激波风洞中瞬态热流测量技术的研究,针对尖化前缘尺度小、热流大的特点,采用传感器-模型一体化技术,研制了适用于无后掠尖化前缘驻点热流测量的整体式薄膜电阻温度计。对经典的Fay-Riddell驻点热流公式进行研究,实验验证了Fay-Riddell公式在球头半径较大时的工程实用性。研究了低雷诺数、稀薄气体效应对驻点区流动的影响,应用了多个适用于尖头体、尖化前缘的驻点热流工程计算关系式,获得了此类外形驻点区域气动加热的预测方法。
在FD-20激波风洞中对尖化前缘外形进行了热环境实验研究。实验模型采用R=1、1.5、3和5mm柱-楔外形的传感器与钢质基座组合模型。通过实验,获得Ma=5、6条件下无后掠尖化前缘驻点热流,数据的重复性误差控制在15%以内。研究表明,在R=3、5mm时,驻点热流测量结果与连续理论预测结果相差较小,经典的边界层理论仍能适用。而随着雷诺数减小,R=1、1.5mm时,由于激波厚度增大,薄边界层将朝着强激波方向发展,驻点区域出现多种流动的干扰,低雷诺数效应出现,热传导开始增大。将实验数据与理论预测结果进行比较,R=1、1.5mm前缘驻点热流测量结果比连续理论预测结果增大10%～20%,与修正边界层理论预测趋势一致。本文还对今后研究提出具体要求。
As the development of new hypersonic vehicles, which can make a long time and long range flight, configuration with high ratio of lift to drag becomes the key point. For the blunt wedges could not meet the needs of air performances of the new hypersonic vehicles, the present of sharp leading edge with minimal drag becomes reasonable. Sharp leading edges offer numerous advantages on air performances. Meanwhile sharp leading edges bring high level of aero heating and difficulty of thermal protection. As the new hypersonic vehicles want to maintain its shape and high ratio of lift to drag, it has to employ non ablating thermal protection, which needs high precision heat-transfer data. It is of great significance for the development of new hypersonic vehicles to make experimental technique study of heat transfer measurement of sharp leading edges, which would provide reliable data for theoretical predictions and thermal protection researches.
Investigations on transient heat transfer measurement in shock tunnel have been done. For the characters of small scale and high heating level of sharp leading edges, an instrument-model integration technique was occupied to develop a kind of integration thin film gauge, which was suitable for measurement of stagnation-point heat transfer on the sharp leading edges. Studies on Fay-Riddell relationship, low Reynolds number theory and rarefied flow effects have been conducted. Through studying on the flow properties of stagnation region of nose tips and sharp leading edges, methods of prediction of aero heating on stagnation point of such configurations have been achieved.
Measurement of heat transfer on sharp leading edges has been carried out in shock tunnel FD-20. In the experiments, several cylinder-wedge geometry models, which were the combinations of gauges and steel bases, were employed. Those models had a range of nose radius from 1 - 5 mm. Results of stagnation-point heat transfer of unswept leading edges were obtained for Mach 5 and 6. The repeatability error of those results was less than 15%. In the radius range of 3 - 5 mm, measurement results of heat transfer on the stagnation point were close to that of continuum theory, indicating that classical boundary layer theory could apply. As the Reynolds numbers decreased, in the radius range of 1 - 1.5 mm, the shock thickness increased, and the thin boundary layer grew out toward the strong shock wave. There exited large flow interference in the stagnation region, effect of low Reynolds numbers appeared, and experimental results were larger than that of continuum theory by 10% to 20%. The increasing trend of experiment results was in agreement with modified boundary layer theory. At last this paper presents requirements for future works.
本文通过对激波风洞中瞬态热流测量技术的研究,针对尖化前缘尺度小、热流大的特点,采用传感器-模型一体化技术,研制了适用于无后掠尖化前缘驻点热流测量的整体式薄膜电阻温度计。对经典的Fay-Riddell驻点热流公式进行研究,实验验证了Fay-Riddell公式在球头半径较大时的工程实用性。研究了低雷诺数、稀薄气体效应对驻点区流动的影响,应用了多个适用于尖头体、尖化前缘的驻点热流工程计算关系式,获得了此类外形驻点区域气动加热的预测方法。
在FD-20激波风洞中对尖化前缘外形进行了热环境实验研究。实验模型采用R=1、1.5、3和5mm柱-楔外形的传感器与钢质基座组合模型。通过实验,获得Ma=5、6条件下无后掠尖化前缘驻点热流,数据的重复性误差控制在15%以内。研究表明,在R=3、5mm时,驻点热流测量结果与连续理论预测结果相差较小,经典的边界层理论仍能适用。而随着雷诺数减小,R=1、1.5mm时,由于激波厚度增大,薄边界层将朝着强激波方向发展,驻点区域出现多种流动的干扰,低雷诺数效应出现,热传导开始增大。将实验数据与理论预测结果进行比较,R=1、1.5mm前缘驻点热流测量结果比连续理论预测结果增大10%～20%,与修正边界层理论预测趋势一致。本文还对今后研究提出具体要求。
As the development of new hypersonic vehicles, which can make a long time and long range flight, configuration with high ratio of lift to drag becomes the key point. For the blunt wedges could not meet the needs of air performances of the new hypersonic vehicles, the present of sharp leading edge with minimal drag becomes reasonable. Sharp leading edges offer numerous advantages on air performances. Meanwhile sharp leading edges bring high level of aero heating and difficulty of thermal protection. As the new hypersonic vehicles want to maintain its shape and high ratio of lift to drag, it has to employ non ablating thermal protection, which needs high precision heat-transfer data. It is of great significance for the development of new hypersonic vehicles to make experimental technique study of heat transfer measurement of sharp leading edges, which would provide reliable data for theoretical predictions and thermal protection researches.
Investigations on transient heat transfer measurement in shock tunnel have been done. For the characters of small scale and high heating level of sharp leading edges, an instrument-model integration technique was occupied to develop a kind of integration thin film gauge, which was suitable for measurement of stagnation-point heat transfer on the sharp leading edges. Studies on Fay-Riddell relationship, low Reynolds number theory and rarefied flow effects have been conducted. Through studying on the flow properties of stagnation region of nose tips and sharp leading edges, methods of prediction of aero heating on stagnation point of such configurations have been achieved.
Measurement of heat transfer on sharp leading edges has been carried out in shock tunnel FD-20. In the experiments, several cylinder-wedge geometry models, which were the combinations of gauges and steel bases, were employed. Those models had a range of nose radius from 1 - 5 mm. Results of stagnation-point heat transfer of unswept leading edges were obtained for Mach 5 and 6. The repeatability error of those results was less than 15%. In the radius range of 3 - 5 mm, measurement results of heat transfer on the stagnation point were close to that of continuum theory, indicating that classical boundary layer theory could apply. As the Reynolds numbers decreased, in the radius range of 1 - 1.5 mm, the shock thickness increased, and the thin boundary layer grew out toward the strong shock wave. There exited large flow interference in the stagnation region, effect of low Reynolds numbers appeared, and experimental results were larger than that of continuum theory by 10% to 20%. The increasing trend of experiment results was in agreement with modified boundary layer theory. At last this paper presents requirements for future works.
引文
[1]姜贵庆,刘连元.高速气流传热与烧蚀热防护.国防工业出版社,2003年.
[2]姜贵庆,张学军,王淑华,艾邦成,俞继军.飞行器尖化前缘的热结构特性.宇航材料工艺,2007年第4期.
[3] Paul Kolodziej, Jefiey D. Bull, Frank S. Milos, et al. Aerothermal performance constraints for small radius leading edges operating at hypervelocity. Ames Research Center, Sep. 1997.
[4] Mayhard L. Hill. Materials for small radius leading edges for hypersonic vehicles. J. Spacecraft. Vol. 5, No. 1, Jan 1968.
[5] Joan Salute, Jeff Bull, Dan Rasky. SHARP-B2: Flight test objectives, project implementation, and initial results. NASA, Jan 2001.
[6] Ingrid Dietlein and Alexander Kopp. System analysis for“Sharp-Edge”re-entry vehicles. AIAA 2009-7427, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, Bremen, Oct 2009.
[7]姜贵庆.非烧蚀热防护与非烧蚀机理.近代空气动力学研讨会文集,2005年.
[8]姜宗林,俞鸿儒.高超声速激波风洞研究进展.力学进展,第39卷第6期,2009年11月.
[9]姜宗林.中国科学院高温气体动力学重点实验室研究进展.力学进展,第40卷第2期,2010年3月.
[10]唐志共主编.高超声速气动力试验.国防工业出版社,2004年.
[11]唐贵明,李静美,俞鸿儒. ?激波风洞中的传热测量结果.中科院力学所,1978年.
[12]王曾亚.薄膜电阻温度计的工艺与标定.全国第一届激波管会议.
[13] Michael S. Holden, Timothy P. Wadhams, Matthew MacLean. Experimental studies in hypersonic flows for facility and code validation,AIAA 2007-1304.
[14] Matthew MacLean, and Michael Holden. Catalytic effects on heat transfer measurements for aerothermal studies with CO2. AIAA 2006-182.
[15] T. P. Wadhams, M. G. MacLean, M. S. Holden, et al. Return to flight testing of a 3.5% scale space shuttle OTS and OT model at Mach numbers of 3.5 and 4.0. AIAA 2006-720.
[16]俞鸿儒,李忠发.测量表面热流率的同轴薄膜热电偶.中科院力学所,1980年3月.
[17] Joseph J. Coblish, Stuart M. Coulter, Joseph D. Norris. Aerothermal measurement improvements using coaxial thermocouples at AEDC hypervelocity wind tunnel No. 9. AIAA 2007-1467.
[18] Scott A. Berry, Robert J. Nowak. Effects on fin leading edge sweep on shock-shock interaction at Mach 6. AIAA-96-0230.
[19]周鸿仁,刘秀蓉,徐蓓娜.直流磁控溅射铂电阻薄膜.电子科技大学学报,第26卷第6期,1997年12月.
[20]麻莳立男著.陈昌存,李兆玉,王普译.薄膜技术基础.电子工业出版社,1988年.
[21]田民波,刘德令编译.薄膜科学与技术手册.机械工业出版社,1991年.
[22]赵世臣.常用金属材料手册.冶金工业出版社,1978年.
[23] J. A. Fay and F. R. Riddell. Theory of stagnation point heat transfer in dissociated air. Journal of the Aeronautical Sciences, Vol. 25 No. 2, Feb, 1958.
[24] Robert D. Quinn, Leslie Gong. A method for calculating transient surface temperature and surface heating rates for high-speed aircraft. NASA/TP-2000-209034, Dec 2000.
[25]杨炳尉,徐居福等译.飞行器气动加热和防热计算手册.《科技简报》编辑组,1978年.
[26]黄志澄.航天空气动力学.北京:宇航出版社,1994年.
[27]瞿章华,刘伟,曾明,柳军.高超声速空气动力学.长沙:国防科技大学出版社,2001年.
[28] Dean A. Kontinos, Ken Gee Dinesh, K. Prabhu. Temperature constraints at the sharp leading edge of a crew transfer vehicle. AIAA 2001-2886.
[29] Antonio Ferri, Victor Zakkay and Lu Ting. Blunt-body heat transfer at hypersonic speed and low Reynolds numbers. J. Aerospace Sci, Vol. 28, 1961.
[30] Antonio Ferri, Victor Zakkay. Measurements of stagnation point heat transfer at low Reynolds numbers. J. Aerospace Sci, Vol. 29, 1962.
[31] Antonio Ferri, Victor Zakkay and Lu Ting . On blunt-body heat transfer at hypersonic speed and low Reynolds numbers. J. Aerospace Sci, Vol. 29, 1962.
[32] Paul Kolodziej. Aerothermal performance constraints for hypervelocity small radius unswept leading edges and nosetips. NASA Technical Memorandum 112204, Jul 1997.
[33] Cheng. H. K. The blunt-body problem in hypersonic flow at low Reynolds number. Cornell Aeronautical Laboratory, CAL Report No. AF-1285-A-10, Jun 1963.
[34] Matting. F. W. Approximate bridging relations in the transitional regime between continuum and free-molecule flows. Journal of Spacecraft and Rockets, Vol. 8, No. 1, pp. 35-40, 1971.
[35]王智慧,鲍麟.具有局部稀薄气体效应的高超声速尖锥气动加热特征研究.计算物理,第27卷第1期,2010年1月.
[36]王智慧,鲍麟,童秉纲.高超声速尖头体驻点热流从连续态过渡到稀薄态的变化特征和桥函数研究.中国科学(G辑:物理学力学天文学),2009年第8期.
[37]王增和.轻活塞炮风洞的运行与校测.七零一所科技报告,1978年.
[38]许小金. FD-20炮风洞参数测量和流场校测及估计平衡活塞重量的讨论.七零一所科技报告,1981年.
[39]沈青.《认识稀薄气体动力学》.力学与实践, 2002年第6期.
[40]张占峰.过渡流区测热风洞实验报告.航天十一院科技报告,2008年.
[41] E. F. Blick. Aerodynamic coefficients in the slip and transition regime. AIAA. J, Vol. l No. 11, 1963.
[42]林建民.爆轰驱动激波风洞中金属成型膜片的探索性试验.第十届全国激波与激波管学术会议论文集,2002年.
[43]陈星.激波风洞双膜段膜片破膜质量的研究.第一届高超声速气动技术交流会论文集,2009年.
[44] D. R. Buttsworth. A finite difference routine for the solution of transient one dimensional heat conduction problems with curvature and temperature-dependent thermal problems. OUEL report number 2130/97.
[2]姜贵庆,张学军,王淑华,艾邦成,俞继军.飞行器尖化前缘的热结构特性.宇航材料工艺,2007年第4期.
[3] Paul Kolodziej, Jefiey D. Bull, Frank S. Milos, et al. Aerothermal performance constraints for small radius leading edges operating at hypervelocity. Ames Research Center, Sep. 1997.
[4] Mayhard L. Hill. Materials for small radius leading edges for hypersonic vehicles. J. Spacecraft. Vol. 5, No. 1, Jan 1968.
[5] Joan Salute, Jeff Bull, Dan Rasky. SHARP-B2: Flight test objectives, project implementation, and initial results. NASA, Jan 2001.
[6] Ingrid Dietlein and Alexander Kopp. System analysis for“Sharp-Edge”re-entry vehicles. AIAA 2009-7427, 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference, Bremen, Oct 2009.
[7]姜贵庆.非烧蚀热防护与非烧蚀机理.近代空气动力学研讨会文集,2005年.
[8]姜宗林,俞鸿儒.高超声速激波风洞研究进展.力学进展,第39卷第6期,2009年11月.
[9]姜宗林.中国科学院高温气体动力学重点实验室研究进展.力学进展,第40卷第2期,2010年3月.
[10]唐志共主编.高超声速气动力试验.国防工业出版社,2004年.
[11]唐贵明,李静美,俞鸿儒. ?激波风洞中的传热测量结果.中科院力学所,1978年.
[12]王曾亚.薄膜电阻温度计的工艺与标定.全国第一届激波管会议.
[13] Michael S. Holden, Timothy P. Wadhams, Matthew MacLean. Experimental studies in hypersonic flows for facility and code validation,AIAA 2007-1304.
[14] Matthew MacLean, and Michael Holden. Catalytic effects on heat transfer measurements for aerothermal studies with CO2. AIAA 2006-182.
[15] T. P. Wadhams, M. G. MacLean, M. S. Holden, et al. Return to flight testing of a 3.5% scale space shuttle OTS and OT model at Mach numbers of 3.5 and 4.0. AIAA 2006-720.
[16]俞鸿儒,李忠发.测量表面热流率的同轴薄膜热电偶.中科院力学所,1980年3月.
[17] Joseph J. Coblish, Stuart M. Coulter, Joseph D. Norris. Aerothermal measurement improvements using coaxial thermocouples at AEDC hypervelocity wind tunnel No. 9. AIAA 2007-1467.
[18] Scott A. Berry, Robert J. Nowak. Effects on fin leading edge sweep on shock-shock interaction at Mach 6. AIAA-96-0230.
[19]周鸿仁,刘秀蓉,徐蓓娜.直流磁控溅射铂电阻薄膜.电子科技大学学报,第26卷第6期,1997年12月.
[20]麻莳立男著.陈昌存,李兆玉,王普译.薄膜技术基础.电子工业出版社,1988年.
[21]田民波,刘德令编译.薄膜科学与技术手册.机械工业出版社,1991年.
[22]赵世臣.常用金属材料手册.冶金工业出版社,1978年.
[23] J. A. Fay and F. R. Riddell. Theory of stagnation point heat transfer in dissociated air. Journal of the Aeronautical Sciences, Vol. 25 No. 2, Feb, 1958.
[24] Robert D. Quinn, Leslie Gong. A method for calculating transient surface temperature and surface heating rates for high-speed aircraft. NASA/TP-2000-209034, Dec 2000.
[25]杨炳尉,徐居福等译.飞行器气动加热和防热计算手册.《科技简报》编辑组,1978年.
[26]黄志澄.航天空气动力学.北京:宇航出版社,1994年.
[27]瞿章华,刘伟,曾明,柳军.高超声速空气动力学.长沙:国防科技大学出版社,2001年.
[28] Dean A. Kontinos, Ken Gee Dinesh, K. Prabhu. Temperature constraints at the sharp leading edge of a crew transfer vehicle. AIAA 2001-2886.
[29] Antonio Ferri, Victor Zakkay and Lu Ting. Blunt-body heat transfer at hypersonic speed and low Reynolds numbers. J. Aerospace Sci, Vol. 28, 1961.
[30] Antonio Ferri, Victor Zakkay. Measurements of stagnation point heat transfer at low Reynolds numbers. J. Aerospace Sci, Vol. 29, 1962.
[31] Antonio Ferri, Victor Zakkay and Lu Ting . On blunt-body heat transfer at hypersonic speed and low Reynolds numbers. J. Aerospace Sci, Vol. 29, 1962.
[32] Paul Kolodziej. Aerothermal performance constraints for hypervelocity small radius unswept leading edges and nosetips. NASA Technical Memorandum 112204, Jul 1997.
[33] Cheng. H. K. The blunt-body problem in hypersonic flow at low Reynolds number. Cornell Aeronautical Laboratory, CAL Report No. AF-1285-A-10, Jun 1963.
[34] Matting. F. W. Approximate bridging relations in the transitional regime between continuum and free-molecule flows. Journal of Spacecraft and Rockets, Vol. 8, No. 1, pp. 35-40, 1971.
[35]王智慧,鲍麟.具有局部稀薄气体效应的高超声速尖锥气动加热特征研究.计算物理,第27卷第1期,2010年1月.
[36]王智慧,鲍麟,童秉纲.高超声速尖头体驻点热流从连续态过渡到稀薄态的变化特征和桥函数研究.中国科学(G辑:物理学力学天文学),2009年第8期.
[37]王增和.轻活塞炮风洞的运行与校测.七零一所科技报告,1978年.
[38]许小金. FD-20炮风洞参数测量和流场校测及估计平衡活塞重量的讨论.七零一所科技报告,1981年.
[39]沈青.《认识稀薄气体动力学》.力学与实践, 2002年第6期.
[40]张占峰.过渡流区测热风洞实验报告.航天十一院科技报告,2008年.
[41] E. F. Blick. Aerodynamic coefficients in the slip and transition regime. AIAA. J, Vol. l No. 11, 1963.
[42]林建民.爆轰驱动激波风洞中金属成型膜片的探索性试验.第十届全国激波与激波管学术会议论文集,2002年.
[43]陈星.激波风洞双膜段膜片破膜质量的研究.第一届高超声速气动技术交流会论文集,2009年.
[44] D. R. Buttsworth. A finite difference routine for the solution of transient one dimensional heat conduction problems with curvature and temperature-dependent thermal problems. OUEL report number 2130/97.