超燃冲压发动机燃烧室起动过程研究
摘要
超燃冲压发动机能够利用大气中的氧气,不需自身携带氧化剂,因而在高超声速飞行领域,其比冲比火箭发动机更高,是高超声速飞行具有广阔前景的推进系统。应用于空天往返,导弹武器和商业运输都将为该领域带来重大革新。
低马赫数下的超燃发动机燃烧室起动(点火)问题是目前该领域的主要难题之一因低马赫数(Ma4)下,来流的总温低(900K),燃烧室内静温低于燃料的自点火温度,必须使用辅助点火措施,才能够实现超燃发动机在低马赫数下的成功起动。而低马赫数下点火难点主要有：(1)点火技术；(2)液态燃料雾化。本文将针对上述两个问题,分别进行超燃发动机低总温条件下的气态乙烯点火技术研究和液态燃料雾化喷嘴研究。
针对超燃发动机燃烧室在低马赫数下的点火问题,在脉冲燃烧风洞设备上进行了系列的试验研究。试验来流参数为总温935K,总压0.8Mpa,马赫数2.1,试验时间约220-250ms；试验内容包括冷流流场特性试验、冷流混合试验、5种不同点火方式下的燃烧室点火试验,发动机贫油极限和富油极限试验,发动机内注油位置优化试验等。
结合数值模拟和光学显示,研究了不同凹槽构型下燃烧室的冷流流场特性。数值模拟和高速纹影测量结果显示,凹槽深度18mm,长深比10.8的燃烧室流道内激波最强,流道内有明显的附面层分离区,冷流阻力最大。长深比10和10.8的凹槽较长深比7的凹槽阻力大接近一倍。燃料以垂直壁面喷射的方式注入燃烧室内,除凹槽内注油位置外,其他位置的燃料穿透深度不到5mm。凹槽上游和凹槽内注入的燃料大部分富集在凹槽内。节流能够显著增加上壁面燃料的穿透深度,对燃料的混合和分布有较大改善(百分比)。对下壁面注入的燃料改善效果不明显。
针对凹槽深度18mm,长深比10.8的发动机构型,研究了五种不同的辅助点火方式：火炬点火器辅助点火、氢气自燃辅助点火、节流辅助点火、引导氢气+火炬点火器辅助点火和节流+引导氢气辅助点火。其中点火器或节流单独辅助乙烯点火没有成功。氢气自燃辅助点火,点火器+引导氢气辅助点火和节流+引导氢气辅助点火三种点火方式能够在较宽范围内实现乙烯的点火,乙烯单独稳定燃烧的时间超过100ms。0.43g/s-12.68g/s的氢气从凹槽内注入燃烧室能够发生自燃,并作为引导火焰点燃乙烯,流量超过17.27g/s或从其他位置注入燃烧室,氢气不会发生自点火燃烧；利用点火器+引导氢气进行辅助点火时,点火参数范围为：点火器的功能为点燃引导氢气,功率不小于20kW,引导氢气最低当量比为0.05左右,注入时间大于50ms,点火时间不短于20ms即能可靠点火；利用节流+引导氢气辅助点火时,节流气流量不小于发动机入口流量的10%,引导氢气注油压力不低于5.0Mpa,节流位置在离隔离段入口875mm处,点火时间约60ms能够可靠点火,节流位置靠前(745mm),氢气注油压力低于5.0Mpa不能成功点火。
燃料从凹槽上游和凹槽内注入,测量得到的发动机贫油熄火极限接近,约为当量比0.08；富油工作极限差别较大,分别为当量比0.327和0.471：凹槽构型对点火方式的适应较为敏感,长深比为7的凹槽只能在点火器+引导氢气一种辅助点火方式下实现乙烯的成功点火；研究了低马赫数下(Ma4)下的富油工作极限,凹槽上游和凹槽内燃料当量比不能超过0.2,主要燃料应分布在凹槽下游附近,离凹槽出口过远,注入燃烧室的燃料难以被凹槽出口的高温燃气点燃,对发动机性能没有贡献。试验研究找到了较为理想的5点注油方式,总当量比达到0.85,燃烧室推力为1401N,隔离段压力扰动位置为352mm。利用一维燃烧室释热规律优化方法,得到了当量比1.0下的4点注油位置分布,通过二维数值模拟验证发现,推力增加到1753N,将0.848当量比下的推力增加了25.1%,但隔离段压力扰动位置97.6mm,抗反压裕度下降。
借助高速摄影手段,研究了不同点火方式下,燃烧室内的点火和火焰传播过程。试验发现,氢气自燃辅助点火方式下,氢气流量过大,大部分氢气会直接突破剪切层,进入主流,不能发生自点火燃烧,因此应该控制氢气流量。节流+引导氢气辅助点火时,节流的时间过长,燃烧室内的压力在乙烯注入后约20ms,扰动至隔离段入口位置超过发动机的工作范围。利用节流方式,在发动机中营造一定的燃烧室背压,当燃烧室燃烧后压力达到该压力水平时,及时关闭节流,燃烧能够维持,过早关闭节流可能导致点火失败,节流时间过长,将导致燃烧室背压过高,超出发动机工作范围。因此在利用节流进行点火时,一方面要严格控制节流流量,另一方面要适时关闭节流。
自主设计和加工了液态燃料充气雾化喷射试验系统,试验测量和比较了相同喷射压力下,纯液态喷射和雾化喷射的液滴雾化效果,试验结果表明,即使气液比低至1.36%,雾化效果仍较为理想,稳定后液滴的SMD值小于60μm。气液比越高,喷嘴直径越小,越有利于降低雾化液滴的直径。试验中,最好的雾化效果为气液比15.15%,雾化后的SMD小于40μm。气液混合雾化能够提升液滴沿喷射方向的速度,增加了液滴的喷射动量,有利于增加注油穿透深度。
The Scramjet (Supersonic Combustion Ramjet) propulsion, utilizing the oxygen of air instead of carrying oxdizer itself, performes higher performance and lower operation cost than that of rocket engine at high flight Mach number, will show its potential in aero-space transportation system, cruise missile or commercial transportation.
The start-up (igniton) problem of the scramjet combustor at low Mach number is one of the toughest problems. When flying at low Mach number (Ma4), the static temperature of supersonic flow in combustor is lower than the temperature of fuel self-ignition. Some addition technique should be imported to facilitate the igntion. Ignition technique and fuel atomization (for liquid fuel) are two parts than can't be avoided when facing the igntion probblem at low Mach number and also the subjects this paper focus on.
Experimental research on ignition technique at low Mach number is performed on directly-connected pulse facility. The experimental inflow parameters are as following:935K of total temperature,0.8Mpa of total pressure, Mach number2.1at the nozzle exit and stable experimental time of200ms-250ms. Experiments covering cold flow experiments, fuel mixing in cold flow, five methods on combustor ignition, lean blowout limit and rich operation limit experiments for scramjet and fuel injection scheme optimization have been performed.
Characteristic of cold flow in combustor has been studied by employing experimental and numerical techniques, results from numerical simulation and high speed schlieren measurement indicates that the shock waves is stronger in the combutor configuration with the cavity which with D(depth)=18mm and L/D(length/depth)=10.8, obvious boundary layer seperation is located on the bottom wall near the ramp of cavity and induces biggest resistance. The coldflow resistance is related to the L/D ratio of cavity, resicatnce of combustor with L/D=7is only half of that with L/D=10.8and10. Fuel injection penetration at wall injection scheme is lower than5mm. Fuel injected from cavity floor and upstream of cavity most distributes in cavity. Air throttle downstream the cavity improves the penetration and mixing efficiency of the fuel injected from top wall, but has little influence on the fuel injected from bottom wall.
Combustor configuration with cavity (D=18,L/D=10.8) is employed for test model. Igniton at Mach4achived with the assistant of self-igniton of pilot hydrogen, igniter combined pilot hydrogen and air throttel combined pilot hydrogen. Pilot hydrogen injected from cavity floor at mass flow rate of0.43g/s to12.68g/s will be self-ignited, and then ignite ehtylene, if the mass flow rate exceedes17.27g/s, self igniton won't occur. The ignition method by igniter and pilot hydrogen is the most reliable igniton technique, with igniter power no less than20kW, hydrogen equivalence ratio higher than0.05, hydrogen injection time longer than50ms and igniton time longer than20ms, ethylene ignition could succeed. When air throttle and pilot hydrogen flame were employed, to reliaze ethylene igniton the throttle should be locateed at875mm from isolator entrance, the mass flow rate of throttle should not be less than10%of engine inflow mass flow rate, and the hydrogen injection pressure should be5.0Mpa at lest.
Lean blow-out limits are similar with fuel injected from cavity floor and uptream the cavity. But rich operation limits showes obvious differences, when fuel injected from cavity floor, the rich operation limit of engine is at eqivalence ratio of0.471, when injected upstream the cavity, the limit is at equivalence0.327. The parameter of cavity have strong effect on igniton performance, experimental results indicate that cavity with L/D=7performes worse igniton characteristic than L/D=10and10.8. fule injection schemes research offered a equivalence ratio distribution rule that fuel injection befor the exit of cavity should be less than equivalence ratio0.2which could avoid unstart of intake, and fuel injected downsteam the cavity should insure that will be ignited, ortherwise there is no comtribution to the engine's performance. A relative optimiazed injection scheme obtained by experimetal method, that ethylene injected from5different locations, achived total equivalence ratio0.85and1401.1N thrust for the combustor.
Igniton and flame propagation processes have beed record by high speed photograph. Records show that when hydrogen injected from cavity floor at high mass flow rate, the majority will enter the core flow directly and no self-igniton will occure. When employing air throttle for igniton, the throttle time should be shut before the pressure in combustor spread to the isolator entrance.
Areated-liquied injection is an efficient injection scheme for liquid fuel aomization. Injection at gas-liquid ratio1.36%could achive SMD smaller than60μm which is much smaller than that of pure liquid injection. Higher gas-liquid ratio and smaller injector diameter is beneficial to reduce droplet diameter. The smallest SMD of40μm achived at gas-liquid ratio15.15%with injector diameter lmm. also the mean velocity of the droplets are increased by the aerated-liquid jet which shows the possibility to increase the injection penetration in supersonic flow.
低马赫数下的超燃发动机燃烧室起动(点火)问题是目前该领域的主要难题之一因低马赫数(Ma4)下,来流的总温低(900K),燃烧室内静温低于燃料的自点火温度,必须使用辅助点火措施,才能够实现超燃发动机在低马赫数下的成功起动。而低马赫数下点火难点主要有：(1)点火技术；(2)液态燃料雾化。本文将针对上述两个问题,分别进行超燃发动机低总温条件下的气态乙烯点火技术研究和液态燃料雾化喷嘴研究。
针对超燃发动机燃烧室在低马赫数下的点火问题,在脉冲燃烧风洞设备上进行了系列的试验研究。试验来流参数为总温935K,总压0.8Mpa,马赫数2.1,试验时间约220-250ms；试验内容包括冷流流场特性试验、冷流混合试验、5种不同点火方式下的燃烧室点火试验,发动机贫油极限和富油极限试验,发动机内注油位置优化试验等。
结合数值模拟和光学显示,研究了不同凹槽构型下燃烧室的冷流流场特性。数值模拟和高速纹影测量结果显示,凹槽深度18mm,长深比10.8的燃烧室流道内激波最强,流道内有明显的附面层分离区,冷流阻力最大。长深比10和10.8的凹槽较长深比7的凹槽阻力大接近一倍。燃料以垂直壁面喷射的方式注入燃烧室内,除凹槽内注油位置外,其他位置的燃料穿透深度不到5mm。凹槽上游和凹槽内注入的燃料大部分富集在凹槽内。节流能够显著增加上壁面燃料的穿透深度,对燃料的混合和分布有较大改善(百分比)。对下壁面注入的燃料改善效果不明显。
针对凹槽深度18mm,长深比10.8的发动机构型,研究了五种不同的辅助点火方式：火炬点火器辅助点火、氢气自燃辅助点火、节流辅助点火、引导氢气+火炬点火器辅助点火和节流+引导氢气辅助点火。其中点火器或节流单独辅助乙烯点火没有成功。氢气自燃辅助点火,点火器+引导氢气辅助点火和节流+引导氢气辅助点火三种点火方式能够在较宽范围内实现乙烯的点火,乙烯单独稳定燃烧的时间超过100ms。0.43g/s-12.68g/s的氢气从凹槽内注入燃烧室能够发生自燃,并作为引导火焰点燃乙烯,流量超过17.27g/s或从其他位置注入燃烧室,氢气不会发生自点火燃烧；利用点火器+引导氢气进行辅助点火时,点火参数范围为：点火器的功能为点燃引导氢气,功率不小于20kW,引导氢气最低当量比为0.05左右,注入时间大于50ms,点火时间不短于20ms即能可靠点火；利用节流+引导氢气辅助点火时,节流气流量不小于发动机入口流量的10%,引导氢气注油压力不低于5.0Mpa,节流位置在离隔离段入口875mm处,点火时间约60ms能够可靠点火,节流位置靠前(745mm),氢气注油压力低于5.0Mpa不能成功点火。
燃料从凹槽上游和凹槽内注入,测量得到的发动机贫油熄火极限接近,约为当量比0.08；富油工作极限差别较大,分别为当量比0.327和0.471：凹槽构型对点火方式的适应较为敏感,长深比为7的凹槽只能在点火器+引导氢气一种辅助点火方式下实现乙烯的成功点火；研究了低马赫数下(Ma4)下的富油工作极限,凹槽上游和凹槽内燃料当量比不能超过0.2,主要燃料应分布在凹槽下游附近,离凹槽出口过远,注入燃烧室的燃料难以被凹槽出口的高温燃气点燃,对发动机性能没有贡献。试验研究找到了较为理想的5点注油方式,总当量比达到0.85,燃烧室推力为1401N,隔离段压力扰动位置为352mm。利用一维燃烧室释热规律优化方法,得到了当量比1.0下的4点注油位置分布,通过二维数值模拟验证发现,推力增加到1753N,将0.848当量比下的推力增加了25.1%,但隔离段压力扰动位置97.6mm,抗反压裕度下降。
借助高速摄影手段,研究了不同点火方式下,燃烧室内的点火和火焰传播过程。试验发现,氢气自燃辅助点火方式下,氢气流量过大,大部分氢气会直接突破剪切层,进入主流,不能发生自点火燃烧,因此应该控制氢气流量。节流+引导氢气辅助点火时,节流的时间过长,燃烧室内的压力在乙烯注入后约20ms,扰动至隔离段入口位置超过发动机的工作范围。利用节流方式,在发动机中营造一定的燃烧室背压,当燃烧室燃烧后压力达到该压力水平时,及时关闭节流,燃烧能够维持,过早关闭节流可能导致点火失败,节流时间过长,将导致燃烧室背压过高,超出发动机工作范围。因此在利用节流进行点火时,一方面要严格控制节流流量,另一方面要适时关闭节流。
自主设计和加工了液态燃料充气雾化喷射试验系统,试验测量和比较了相同喷射压力下,纯液态喷射和雾化喷射的液滴雾化效果,试验结果表明,即使气液比低至1.36%,雾化效果仍较为理想,稳定后液滴的SMD值小于60μm。气液比越高,喷嘴直径越小,越有利于降低雾化液滴的直径。试验中,最好的雾化效果为气液比15.15%,雾化后的SMD小于40μm。气液混合雾化能够提升液滴沿喷射方向的速度,增加了液滴的喷射动量,有利于增加注油穿透深度。
The Scramjet (Supersonic Combustion Ramjet) propulsion, utilizing the oxygen of air instead of carrying oxdizer itself, performes higher performance and lower operation cost than that of rocket engine at high flight Mach number, will show its potential in aero-space transportation system, cruise missile or commercial transportation.
The start-up (igniton) problem of the scramjet combustor at low Mach number is one of the toughest problems. When flying at low Mach number (Ma4), the static temperature of supersonic flow in combustor is lower than the temperature of fuel self-ignition. Some addition technique should be imported to facilitate the igntion. Ignition technique and fuel atomization (for liquid fuel) are two parts than can't be avoided when facing the igntion probblem at low Mach number and also the subjects this paper focus on.
Experimental research on ignition technique at low Mach number is performed on directly-connected pulse facility. The experimental inflow parameters are as following:935K of total temperature,0.8Mpa of total pressure, Mach number2.1at the nozzle exit and stable experimental time of200ms-250ms. Experiments covering cold flow experiments, fuel mixing in cold flow, five methods on combustor ignition, lean blowout limit and rich operation limit experiments for scramjet and fuel injection scheme optimization have been performed.
Characteristic of cold flow in combustor has been studied by employing experimental and numerical techniques, results from numerical simulation and high speed schlieren measurement indicates that the shock waves is stronger in the combutor configuration with the cavity which with D(depth)=18mm and L/D(length/depth)=10.8, obvious boundary layer seperation is located on the bottom wall near the ramp of cavity and induces biggest resistance. The coldflow resistance is related to the L/D ratio of cavity, resicatnce of combustor with L/D=7is only half of that with L/D=10.8and10. Fuel injection penetration at wall injection scheme is lower than5mm. Fuel injected from cavity floor and upstream of cavity most distributes in cavity. Air throttle downstream the cavity improves the penetration and mixing efficiency of the fuel injected from top wall, but has little influence on the fuel injected from bottom wall.
Combustor configuration with cavity (D=18,L/D=10.8) is employed for test model. Igniton at Mach4achived with the assistant of self-igniton of pilot hydrogen, igniter combined pilot hydrogen and air throttel combined pilot hydrogen. Pilot hydrogen injected from cavity floor at mass flow rate of0.43g/s to12.68g/s will be self-ignited, and then ignite ehtylene, if the mass flow rate exceedes17.27g/s, self igniton won't occur. The ignition method by igniter and pilot hydrogen is the most reliable igniton technique, with igniter power no less than20kW, hydrogen equivalence ratio higher than0.05, hydrogen injection time longer than50ms and igniton time longer than20ms, ethylene ignition could succeed. When air throttle and pilot hydrogen flame were employed, to reliaze ethylene igniton the throttle should be locateed at875mm from isolator entrance, the mass flow rate of throttle should not be less than10%of engine inflow mass flow rate, and the hydrogen injection pressure should be5.0Mpa at lest.
Lean blow-out limits are similar with fuel injected from cavity floor and uptream the cavity. But rich operation limits showes obvious differences, when fuel injected from cavity floor, the rich operation limit of engine is at eqivalence ratio of0.471, when injected upstream the cavity, the limit is at equivalence0.327. The parameter of cavity have strong effect on igniton performance, experimental results indicate that cavity with L/D=7performes worse igniton characteristic than L/D=10and10.8. fule injection schemes research offered a equivalence ratio distribution rule that fuel injection befor the exit of cavity should be less than equivalence ratio0.2which could avoid unstart of intake, and fuel injected downsteam the cavity should insure that will be ignited, ortherwise there is no comtribution to the engine's performance. A relative optimiazed injection scheme obtained by experimetal method, that ethylene injected from5different locations, achived total equivalence ratio0.85and1401.1N thrust for the combustor.
Igniton and flame propagation processes have beed record by high speed photograph. Records show that when hydrogen injected from cavity floor at high mass flow rate, the majority will enter the core flow directly and no self-igniton will occure. When employing air throttle for igniton, the throttle time should be shut before the pressure in combustor spread to the isolator entrance.
Areated-liquied injection is an efficient injection scheme for liquid fuel aomization. Injection at gas-liquid ratio1.36%could achive SMD smaller than60μm which is much smaller than that of pure liquid injection. Higher gas-liquid ratio and smaller injector diameter is beneficial to reduce droplet diameter. The smallest SMD of40μm achived at gas-liquid ratio15.15%with injector diameter lmm. also the mean velocity of the droplets are increased by the aerated-liquid jet which shows the possibility to increase the injection penetration in supersonic flow.
引文
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