航空燃气轮机涡轮级间燃烧技术研究
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摘要
航空燃气轮机性能的提高,需要在提高总增压比的同时,进一步提高涡轮进口温度,但现有技术条件下,压气机总增压比越高,其喘振裕度越小,而涡轮进口总温也受到耐热材料的限制,这都限制着航空发动机技术的发展。在高低压涡轮涡轮之间设置级间燃烧室,通过再热循环提高发动机总体性能被认为是现有技术条件下,增大发动机单位推力的一个有效手段。
在总结当今航空发动机涡轮级间燃烧技术的基础上,分析了现有研究中存在的问题,采用级间燃烧室后的发动机总体性能计算上存在两种不同的观点,一种观点认为耗油率近似不变即可显著提高发动机单位推力,一种认为单位推力的提高需要相近增量的耗油率,而涡轮级间燃烧室试验中出口温度场也难以匹配涡轮耐热需求,是现有研究中遇到的两个主要问题。
针对总体性能计算结果的观点分歧问题,本文在非理想循环下编制了带级间燃烧室的航空发动机总体性能计算程序,对比了部分理想循环与非理想循环的总体性能差异,采用加力比、耗油率比等指标分析了级间燃烧室的适用范围,研究了高温升燃烧室发动机与带级间燃烧室发动机的各自优、缺点,指出了主燃烧室与级间燃烧室的热量分配方法,最后进行了带级间燃烧室发动机的优化计算,通过对F100-PW229的计算显示,优化后的发动机耗油率增长7%,单位推力增加30%。
对某型发动机的过渡段设置级间燃烧室的方案进行了数值模拟研究,研究了过渡段外围周向设置凹腔、叶片径向开凹腔的方案,并在上述方案的基础上在周向凹腔引入冷却气进行了研究,还研究了采用环向的v型稳定器燃烧方案以及多个径向稳定器方案,径向稳定器方案中又研究了单点喷射燃料和多点喷射燃料,最终选取采用多个径向稳定器、在叶片尾缘多点喷射燃料作为下一步试验的方案。
为开展涡轮流道内的液体燃料喷射特性试验,采用Fraunhofer衍射技术和CCD相机分别测量了涡轮导叶流道内的液滴雾化细度和射流轨迹,试验来流速度56m/s-91m/s,喷嘴压差50kPa～400kPa,在叶片前缘喷射燃料,分别向吸力面、压力面侧喷和流道中心逆喷,研究发现：涡轮叶片流道内液滴轨迹同时受液气动量比、叶片折转角度、稠度的影响,在向吸力面侧喷时,液滴距离叶背太近,容易灼伤叶背,在向压力面侧喷时,液滴会打到叶片尾缘上,容易形成积碳,在中心逆喷时,液滴轨迹距离叶片表面相对较远。涡轮流道内液滴横向喷射的雾化粒径可以满足燃烧室要求。
在之前研究的基础上,采用径向稳定器进行了燃烧实验,考察了点、熄火特性、燃烧效率和出口温度场,喷油杆结构分为单孔喷射和三孔喷射,V型稳定器尾缘宽度分别为9mm和13mm,共在0.2kg/s、0.25kg/s、0.3kg/s、0.35kg/s/、0.4kg/s五个来流空气流量下进行了试验,来流温度为573K,试验结果显示：采用三孔喷油杆顺喷时的点熄火特性最好,稳定器尾缘宽度13mm,斜侧喷时的燃烧效率最高,温度场最均匀。
下一步可在过渡段设置全径向稳定器,进行真实工况的环形叶栅燃烧试验。
In order to improve performance of aviation gas turbine, the turbine inlet temperature must be increased while enhancing the compressor over pressure ratio, but under modern technical conditions, the compressor over pressure ratio was confined by its surge margin, and turbine materials limited the turbine inlet temperature, these above conditions limited the development of aero engine technology. Adding a inter-stage turbine burner between high pressure turbine and low pressure turbine was seemed a effective method to increase specific thrust.
Summarizing today's gas turbine engine inter-stage turbine burner technology, analyzing the existing problem in nowadays research, results show that there are two different views on the engine overall performance calculations, one view is that inter-stage turbine burner significantly enhancing specific thrust while increasing little thrust specific fuel consumption, the other view is that specific thrust increasing need similar incremental fuel consumption, and non-uniform outlet temperature distribution of combustor is also an important problem in the experiment.
Compiling a program to resolve the opinions differences of engine overall performance, comparing the different performances between partially ideal cycle and non-ideal cycle, using index such as specific thrust increase ratio and fuel consumption increase ratio, to analyze flight target of inter-stage turbine burner engine, researching the advantages and disadvantages of high temperature rise combustor engine and inter-stage turbine burner engine, point out heat distribution methods of the main burner and secondary combustor. Finally the inter-stage turbine burner engine overall performance was optimized calculation, specific thrust increase30%while fuel consumption increase7%under F100-PW229parameters.
Four cases of adding a inter-stage turbine burner in the engine inter transition duct were investigated by numerical simulation, firstly, the case of setting up a circumferential cavity outside the inter transition duct and a radial cavity on the vane was studied, secondly, based on the first case, cooling gas injecting into circumferential cavity was calculated, thirdly, a case of installing a circumferential V-gutter flame stabilizer in the duct was researched, fourthly, a case of setting up several radical V-gutter flame stabilizer while injecting fuel from a single injector and three injectors was studied, finally, the forth case with three injectors was selected as the experimental case in research program.
Liquid fuel spray characteristics in blade passage of low pressure turbine was tested, Fraunhofer diffraction technique was used to measure droplet diameter, and the spray trajectory was recorded by CCD camera, the inlet velocity in experiments is56m/s～91m/s and the pressure differential of injector is50kPa-400kPa, fuel was injected to pressure side, suction side and reverse side at the blade leading edge respectively. Results show that spray trajectory in turbine passage is influenced by liquid-air momentum ratio, blade-turning angle as well as blade solidity. Spray trajectory is close to blade surface when injecting to blade passage, and flame is easy to burn out the vane, and droplet will hit the blade trailing edge to form carbon deposit when injecting to pressure side, droplet is far from the blade when reverse injection at the middle of passage. At real working condition, the spray fineness can meet combustor requirement.
Based on above research, a combustion experiment with radical V-gutter flame stabilizer was tested, the lean blowout, lean light-up characteristics, combustion efficiency and outlet temperature distribution were evaluated. A single hole injector and three-hole injector were used to spray, the trailing edge width of V-gutter was respectively9mm and13mm, the inlet mass flow rate is0.2kg/s,0.25kg/s,0.3kg/s,0.35kg/s/,0.4kg/s, and the temperature of inlet flow is573K. Results show that the three-hole injector spraying along the flow has a better ignition and flameout feature, the trailing edge width of the stabilizer of13mm, lateral and oblique injection has the highest combustion efficiency, and also have a well-distributed temperature.
An experiment of installing several radical V-gutter stabilizers under real condition should be tested in the future.
在总结当今航空发动机涡轮级间燃烧技术的基础上,分析了现有研究中存在的问题,采用级间燃烧室后的发动机总体性能计算上存在两种不同的观点,一种观点认为耗油率近似不变即可显著提高发动机单位推力,一种认为单位推力的提高需要相近增量的耗油率,而涡轮级间燃烧室试验中出口温度场也难以匹配涡轮耐热需求,是现有研究中遇到的两个主要问题。
针对总体性能计算结果的观点分歧问题,本文在非理想循环下编制了带级间燃烧室的航空发动机总体性能计算程序,对比了部分理想循环与非理想循环的总体性能差异,采用加力比、耗油率比等指标分析了级间燃烧室的适用范围,研究了高温升燃烧室发动机与带级间燃烧室发动机的各自优、缺点,指出了主燃烧室与级间燃烧室的热量分配方法,最后进行了带级间燃烧室发动机的优化计算,通过对F100-PW229的计算显示,优化后的发动机耗油率增长7%,单位推力增加30%。
对某型发动机的过渡段设置级间燃烧室的方案进行了数值模拟研究,研究了过渡段外围周向设置凹腔、叶片径向开凹腔的方案,并在上述方案的基础上在周向凹腔引入冷却气进行了研究,还研究了采用环向的v型稳定器燃烧方案以及多个径向稳定器方案,径向稳定器方案中又研究了单点喷射燃料和多点喷射燃料,最终选取采用多个径向稳定器、在叶片尾缘多点喷射燃料作为下一步试验的方案。
为开展涡轮流道内的液体燃料喷射特性试验,采用Fraunhofer衍射技术和CCD相机分别测量了涡轮导叶流道内的液滴雾化细度和射流轨迹,试验来流速度56m/s-91m/s,喷嘴压差50kPa～400kPa,在叶片前缘喷射燃料,分别向吸力面、压力面侧喷和流道中心逆喷,研究发现：涡轮叶片流道内液滴轨迹同时受液气动量比、叶片折转角度、稠度的影响,在向吸力面侧喷时,液滴距离叶背太近,容易灼伤叶背,在向压力面侧喷时,液滴会打到叶片尾缘上,容易形成积碳,在中心逆喷时,液滴轨迹距离叶片表面相对较远。涡轮流道内液滴横向喷射的雾化粒径可以满足燃烧室要求。
在之前研究的基础上,采用径向稳定器进行了燃烧实验,考察了点、熄火特性、燃烧效率和出口温度场,喷油杆结构分为单孔喷射和三孔喷射,V型稳定器尾缘宽度分别为9mm和13mm,共在0.2kg/s、0.25kg/s、0.3kg/s、0.35kg/s/、0.4kg/s五个来流空气流量下进行了试验,来流温度为573K,试验结果显示：采用三孔喷油杆顺喷时的点熄火特性最好,稳定器尾缘宽度13mm,斜侧喷时的燃烧效率最高,温度场最均匀。
下一步可在过渡段设置全径向稳定器,进行真实工况的环形叶栅燃烧试验。
In order to improve performance of aviation gas turbine, the turbine inlet temperature must be increased while enhancing the compressor over pressure ratio, but under modern technical conditions, the compressor over pressure ratio was confined by its surge margin, and turbine materials limited the turbine inlet temperature, these above conditions limited the development of aero engine technology. Adding a inter-stage turbine burner between high pressure turbine and low pressure turbine was seemed a effective method to increase specific thrust.
Summarizing today's gas turbine engine inter-stage turbine burner technology, analyzing the existing problem in nowadays research, results show that there are two different views on the engine overall performance calculations, one view is that inter-stage turbine burner significantly enhancing specific thrust while increasing little thrust specific fuel consumption, the other view is that specific thrust increasing need similar incremental fuel consumption, and non-uniform outlet temperature distribution of combustor is also an important problem in the experiment.
Compiling a program to resolve the opinions differences of engine overall performance, comparing the different performances between partially ideal cycle and non-ideal cycle, using index such as specific thrust increase ratio and fuel consumption increase ratio, to analyze flight target of inter-stage turbine burner engine, researching the advantages and disadvantages of high temperature rise combustor engine and inter-stage turbine burner engine, point out heat distribution methods of the main burner and secondary combustor. Finally the inter-stage turbine burner engine overall performance was optimized calculation, specific thrust increase30%while fuel consumption increase7%under F100-PW229parameters.
Four cases of adding a inter-stage turbine burner in the engine inter transition duct were investigated by numerical simulation, firstly, the case of setting up a circumferential cavity outside the inter transition duct and a radial cavity on the vane was studied, secondly, based on the first case, cooling gas injecting into circumferential cavity was calculated, thirdly, a case of installing a circumferential V-gutter flame stabilizer in the duct was researched, fourthly, a case of setting up several radical V-gutter flame stabilizer while injecting fuel from a single injector and three injectors was studied, finally, the forth case with three injectors was selected as the experimental case in research program.
Liquid fuel spray characteristics in blade passage of low pressure turbine was tested, Fraunhofer diffraction technique was used to measure droplet diameter, and the spray trajectory was recorded by CCD camera, the inlet velocity in experiments is56m/s～91m/s and the pressure differential of injector is50kPa-400kPa, fuel was injected to pressure side, suction side and reverse side at the blade leading edge respectively. Results show that spray trajectory in turbine passage is influenced by liquid-air momentum ratio, blade-turning angle as well as blade solidity. Spray trajectory is close to blade surface when injecting to blade passage, and flame is easy to burn out the vane, and droplet will hit the blade trailing edge to form carbon deposit when injecting to pressure side, droplet is far from the blade when reverse injection at the middle of passage. At real working condition, the spray fineness can meet combustor requirement.
Based on above research, a combustion experiment with radical V-gutter flame stabilizer was tested, the lean blowout, lean light-up characteristics, combustion efficiency and outlet temperature distribution were evaluated. A single hole injector and three-hole injector were used to spray, the trailing edge width of V-gutter was respectively9mm and13mm, the inlet mass flow rate is0.2kg/s,0.25kg/s,0.3kg/s,0.35kg/s/,0.4kg/s, and the temperature of inlet flow is573K. Results show that the three-hole injector spraying along the flow has a better ignition and flameout feature, the trailing edge width of the stabilizer of13mm, lateral and oblique injection has the highest combustion efficiency, and also have a well-distributed temperature.
An experiment of installing several radical V-gutter stabilizers under real condition should be tested in the future.
引文
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