光壁面超声速燃烧室点火及火焰稳定研究
摘要
超燃冲压发动机作为大气层内高超声速飞行的最佳动力装置，超声速燃烧室的综合性能至关重要。光壁面超声速燃烧室壁面没有凹槽等高热流部件，有利于冷却流道设计，中心燃烧组织方式进一步降低了平均热流。燃烧稳定过程中采用的支板能够增强燃料的掺混，缩短燃烧距离，降低燃烧室的长度。已研究的光壁面超声速燃烧室模型一般都采用氢气作为燃料，而真正飞行器在飞行马赫数Ma=4-7范围内一般都采用液态航空煤油，其反应时间的延长会使得光壁面超声速燃烧室内的燃烧组织面临较大的困难。本文致力于解决液态煤油在光壁面超声速燃烧室的点火及火焰稳定问题，主要研究内容如下：
首先设计制作了一种小型高焓气流发生器并将之安装在光壁面超声速燃烧室的壁面上，利用高焓气流发生器形成的常驻火焰来点火和维持燃烧稳定。点火过程中借助了空气节流，通过改变节流空气压力和节流空气的作用时间来控制点火过程中的节流强度。研究发现不同的节流强度会在燃烧室内形成四种不同的点火模式：稳定模式、不起动模式、不稳定模式和熄火模式。研究分析了不同当量比下点火模式的分布规律，给出了基于这种燃烧组织方式燃烧室的稳定点火边界。研究结果证明利用常驻火焰可以在光壁面超声速燃烧室内实现煤油的稳定燃烧。
进一步的研究中将小型高焓气流发生器镶嵌在了喷油支板内形成火箭支板，直接在支板尾部形成常驻火焰。实验测试了两种火箭羽流条件下的点火结果：一种条件下火箭支板喷射高温燃气，另一种条件下火箭支板只喷射氧气。研究结果发现当火箭支板只喷射氧气时更适合光壁面超声速燃烧室的点火和燃烧稳定。基于此设计出了补氧支板结构，通过在支板尾部直接注入氧气，将支板尾部低速区拉长，加快燃料反应速度，成功地在高速冷气流中形成了稳定的火焰。研究发现采用补氧支板时，支板尾部火焰的存在并不意味着燃烧室整体燃烧的建立，整体燃烧需要氧气流量达到一定程度才能够触发。
进一步研究了基于补氧支板的光壁面超声速燃烧室燃烧组织特性，发现采用补氧支板，大大拓宽了燃烧室的燃烧稳定边界，在当量比ER=0.19的条件下依然能够保证燃烧的稳定，并且点火冲击小。随着当量比的增加，燃烧过程会出现热阻塞，燃烧室内会出现亚声速区。研究中根据燃烧室内亚声速区所占比例的不同，将燃烧室内的燃烧模态分为三种：超燃模态、亚燃模态和过渡模态。研究发现不同燃烧模态下的火焰传播过程截然不同，当燃烧室出现热阻塞时，整体燃烧建立过程有明显的阶段性，火焰传播方向与燃烧室气体流动方向相反，而在超燃状态下整体燃烧火焰快速建立，不存在火焰逆流传播现象。
研究发现当光壁面超声速燃烧室的整体燃烧建立以后，维持整体燃烧所需要的氧气量要明显小于整体燃烧建立所需要的氧气量，并且随着当量比的增加，维持燃烧所需要的氧气量变得越来越少。基于此规律，研究给出了燃烧过程中氧气供应方案的优化思路：点火阶段采取大氧气流量供应，而燃烧稳定阶段则采取小氧气流量供应，这样有助于降低飞行器在实际飞行过程中氧气的携带量。文章研究了支板上不同燃料喷射集中度对燃烧室压力分布的影响，发现改变支板上的燃料喷射集中度主要影响支板尾部附近区域的压升，对燃烧室尾部的影响不大。
最后，为了进一步降低燃烧室热防护压力，提出了一种支板/壁面组合燃料喷射方法。初步研究发现将支板喷射的煤油转移到壁面喷射之后，燃烧特性会受到影响。但是通过优化支板/壁面燃料喷射比例，在保证燃烧效果相近的前提下，能够使燃烧室壁面的温度在一定区域内显著降低。研究表明这种燃料喷注方式还能够满足宽马赫数下多点燃料喷射的需求，壁面注入的燃料能够很好地参与到燃烧反应中去，但需注意燃料的喷射位置和喷射方向。
Scramjet is the best propulsion device for the hypersonic vehicle within the atmosphere and the overall performance of its key component combustor is very important. A flush wall combustor is usually take the strut as the flame holder, the combustion start in the center of the combustor, avoid direct contact with the combustion zone, heat load of the wall is reduced. Meanwhile the flush wall is favorable to the cooling channel design. Due to the enhancement of the strut fuel injection, the burning length can be shortened which means a relative shorter aircraft length. All these advantages help to improve the scramjet performance. But most research related to the flush wall supersonic combustor take the hydrogen as the fuel, if using kerosene, the relative long reacting time will greatly increase the difficulty to the combustion organization. This paper has tried to solve this problem and the main contents are as follows:
First, a Micro Hot Gas Generator (MHGG) is designed and mounted on the combustor wall and combustion is to be realized by this permanent flame generated by this MHGG. During the ignition process, air throttling is introduced for ignition auxiliary. By changing the throttling air pressure and acting time, different throttling intensity can be controlled and led to four ignition modes. Too intense of the throttling will cause a un-start phenomenon while too weak of the throttling cannot get stable ignition. The law of the ignition mode distribution under different ER and the stable ignition boundary are gotten via analyzing a large number experiment data. This research has proved the feasibility of stable combustion in a flush wall supersonic combustor by the permanent flame.
Further research has combined the MHGG with the fuel injection strut forming a strut-jet structure, based on this structure, permanent flame is moved the strut back. Two kinds of jet flow are tested: on is high temperature burned gas and the other is pure oxygen. It is discovered that the pure-oxygen plume is more suitable for the global ignition in the flush wall combustor and an oxygen pilot combustion strategy is then developed. Based on this, an oxygen-pilot strut is proposed. With the help of oxygen injection, low speed zone at the strut back is enlarged and the reaction speed is increased, then stable local flame is realized at the back of the strut in the high speed cold air flow. Based on this oxygen pilot strut, local flame at the strut back is not mean global combustion in the flush wall supersonic combustor; global combustion cannot be triggered until the oxygen mass flow rate reaches a certain level.
The paper has investigated the organization characteristics, based on this oxygen pilot strut, stable combustion can be realized in a wide range of ER in the flush wall supersonic combustor, even at a low ER=0.19. Mean while, the ignition process is much steadier compared with the air throttling way. With the increase of ER, thermal choking will occur which means subsonic combustion zone in the combustor. Base on the proportion of the subsonic zone combustion modes can be divided to three kinds: supersonic combustion mode, dual combustion mode and subsonic combustion mode. It is discovered that, under different combustion mode, the flame transition process is totally different. When thermal choking occurs in the combustor, the global combustion is developed from back to front, this counter flow flame transition phenomenon does not happen in the supersonic mode.
When global combustion is founded, the oxygen mass flow rate need to sustain the combustion is much lower than needed during the ignition process; based on this discovery, the optimization program of oxygen supply during the combustion process is given: large oxygen flow rate during the ignition stage and low oxygen flow rate during the stable combustion stage. This can reduce the load of the aircraft during the real flight process. The paper also investigated the influence of the fuel injection concentration to the combustion process. It is discovered the main area affected by the fuel injection concentration is in the strut nearby area, not so obvious at the rear part of the combustor.
Finally a strut/wall combined fuel injection strategy is proposed. Experiment results show that, combustion process in the flush wall supersonic combustor will be affected if the fuel from the strut is moved to the wall. However, by optimizing the fuel proportion between the strut and the wall, the wall temperature can be noticeable reduced without affecting the combustion process seriously. Meanwhile, this strut/wall combined fuel injection strategy can also meet the multi-points fuel injection requirement at the other flight Mach number. To guaranty the combustion performance, the wall injection should pay attention to the injection location and direction.
首先设计制作了一种小型高焓气流发生器并将之安装在光壁面超声速燃烧室的壁面上，利用高焓气流发生器形成的常驻火焰来点火和维持燃烧稳定。点火过程中借助了空气节流，通过改变节流空气压力和节流空气的作用时间来控制点火过程中的节流强度。研究发现不同的节流强度会在燃烧室内形成四种不同的点火模式：稳定模式、不起动模式、不稳定模式和熄火模式。研究分析了不同当量比下点火模式的分布规律，给出了基于这种燃烧组织方式燃烧室的稳定点火边界。研究结果证明利用常驻火焰可以在光壁面超声速燃烧室内实现煤油的稳定燃烧。
进一步的研究中将小型高焓气流发生器镶嵌在了喷油支板内形成火箭支板，直接在支板尾部形成常驻火焰。实验测试了两种火箭羽流条件下的点火结果：一种条件下火箭支板喷射高温燃气，另一种条件下火箭支板只喷射氧气。研究结果发现当火箭支板只喷射氧气时更适合光壁面超声速燃烧室的点火和燃烧稳定。基于此设计出了补氧支板结构，通过在支板尾部直接注入氧气，将支板尾部低速区拉长，加快燃料反应速度，成功地在高速冷气流中形成了稳定的火焰。研究发现采用补氧支板时，支板尾部火焰的存在并不意味着燃烧室整体燃烧的建立，整体燃烧需要氧气流量达到一定程度才能够触发。
进一步研究了基于补氧支板的光壁面超声速燃烧室燃烧组织特性，发现采用补氧支板，大大拓宽了燃烧室的燃烧稳定边界，在当量比ER=0.19的条件下依然能够保证燃烧的稳定，并且点火冲击小。随着当量比的增加，燃烧过程会出现热阻塞，燃烧室内会出现亚声速区。研究中根据燃烧室内亚声速区所占比例的不同，将燃烧室内的燃烧模态分为三种：超燃模态、亚燃模态和过渡模态。研究发现不同燃烧模态下的火焰传播过程截然不同，当燃烧室出现热阻塞时，整体燃烧建立过程有明显的阶段性，火焰传播方向与燃烧室气体流动方向相反，而在超燃状态下整体燃烧火焰快速建立，不存在火焰逆流传播现象。
研究发现当光壁面超声速燃烧室的整体燃烧建立以后，维持整体燃烧所需要的氧气量要明显小于整体燃烧建立所需要的氧气量，并且随着当量比的增加，维持燃烧所需要的氧气量变得越来越少。基于此规律，研究给出了燃烧过程中氧气供应方案的优化思路：点火阶段采取大氧气流量供应，而燃烧稳定阶段则采取小氧气流量供应，这样有助于降低飞行器在实际飞行过程中氧气的携带量。文章研究了支板上不同燃料喷射集中度对燃烧室压力分布的影响，发现改变支板上的燃料喷射集中度主要影响支板尾部附近区域的压升，对燃烧室尾部的影响不大。
最后，为了进一步降低燃烧室热防护压力，提出了一种支板/壁面组合燃料喷射方法。初步研究发现将支板喷射的煤油转移到壁面喷射之后，燃烧特性会受到影响。但是通过优化支板/壁面燃料喷射比例，在保证燃烧效果相近的前提下，能够使燃烧室壁面的温度在一定区域内显著降低。研究表明这种燃料喷注方式还能够满足宽马赫数下多点燃料喷射的需求，壁面注入的燃料能够很好地参与到燃烧反应中去，但需注意燃料的喷射位置和喷射方向。
Scramjet is the best propulsion device for the hypersonic vehicle within the atmosphere and the overall performance of its key component combustor is very important. A flush wall combustor is usually take the strut as the flame holder, the combustion start in the center of the combustor, avoid direct contact with the combustion zone, heat load of the wall is reduced. Meanwhile the flush wall is favorable to the cooling channel design. Due to the enhancement of the strut fuel injection, the burning length can be shortened which means a relative shorter aircraft length. All these advantages help to improve the scramjet performance. But most research related to the flush wall supersonic combustor take the hydrogen as the fuel, if using kerosene, the relative long reacting time will greatly increase the difficulty to the combustion organization. This paper has tried to solve this problem and the main contents are as follows:
First, a Micro Hot Gas Generator (MHGG) is designed and mounted on the combustor wall and combustion is to be realized by this permanent flame generated by this MHGG. During the ignition process, air throttling is introduced for ignition auxiliary. By changing the throttling air pressure and acting time, different throttling intensity can be controlled and led to four ignition modes. Too intense of the throttling will cause a un-start phenomenon while too weak of the throttling cannot get stable ignition. The law of the ignition mode distribution under different ER and the stable ignition boundary are gotten via analyzing a large number experiment data. This research has proved the feasibility of stable combustion in a flush wall supersonic combustor by the permanent flame.
Further research has combined the MHGG with the fuel injection strut forming a strut-jet structure, based on this structure, permanent flame is moved the strut back. Two kinds of jet flow are tested: on is high temperature burned gas and the other is pure oxygen. It is discovered that the pure-oxygen plume is more suitable for the global ignition in the flush wall combustor and an oxygen pilot combustion strategy is then developed. Based on this, an oxygen-pilot strut is proposed. With the help of oxygen injection, low speed zone at the strut back is enlarged and the reaction speed is increased, then stable local flame is realized at the back of the strut in the high speed cold air flow. Based on this oxygen pilot strut, local flame at the strut back is not mean global combustion in the flush wall supersonic combustor; global combustion cannot be triggered until the oxygen mass flow rate reaches a certain level.
The paper has investigated the organization characteristics, based on this oxygen pilot strut, stable combustion can be realized in a wide range of ER in the flush wall supersonic combustor, even at a low ER=0.19. Mean while, the ignition process is much steadier compared with the air throttling way. With the increase of ER, thermal choking will occur which means subsonic combustion zone in the combustor. Base on the proportion of the subsonic zone combustion modes can be divided to three kinds: supersonic combustion mode, dual combustion mode and subsonic combustion mode. It is discovered that, under different combustion mode, the flame transition process is totally different. When thermal choking occurs in the combustor, the global combustion is developed from back to front, this counter flow flame transition phenomenon does not happen in the supersonic mode.
When global combustion is founded, the oxygen mass flow rate need to sustain the combustion is much lower than needed during the ignition process; based on this discovery, the optimization program of oxygen supply during the combustion process is given: large oxygen flow rate during the ignition stage and low oxygen flow rate during the stable combustion stage. This can reduce the load of the aircraft during the real flight process. The paper also investigated the influence of the fuel injection concentration to the combustion process. It is discovered the main area affected by the fuel injection concentration is in the strut nearby area, not so obvious at the rear part of the combustor.
Finally a strut/wall combined fuel injection strategy is proposed. Experiment results show that, combustion process in the flush wall supersonic combustor will be affected if the fuel from the strut is moved to the wall. However, by optimizing the fuel proportion between the strut and the wall, the wall temperature can be noticeable reduced without affecting the combustion process seriously. Meanwhile, this strut/wall combined fuel injection strategy can also meet the multi-points fuel injection requirement at the other flight Mach number. To guaranty the combustion performance, the wall injection should pay attention to the injection location and direction.
引文
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