怎么下载tp钱包|rbcc
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怎么看西北工业大学牵头研制的火箭基组合循环发动机“飞天一号”发射成功？ - 知乎
怎么看西北工业大学牵头研制的火箭基组合循环发动机“飞天一号”发射成功？ - 知乎首页知乎知学堂发现等你来答切换模式登录/注册西北工业大学超燃冲压发动机高超声速武器飞天一号怎么看西北工业大学牵头研制的火箭基组合循环发动机“飞天一号”发射成功？火箭基组合循环发动机是可以实现从零启动加速至高超声速的动力系统，研制难度比超燃冲压发动机更大，是未来实现单级入轨，空间运输的关键动力设施。 [图片]显示全部 关注者33被浏览21,992关注问题写回答邀请回答好问题添加评论分享5 个回答默认排序Rocket Docter三流大学在读博士，为中华民族伟大复兴而读书 关注我国RBCC的关键技术更进一步，值得庆祝。国际上两类主流的组合动力发动机就火箭基组合动力循环发动机（rocket-based combined-cycle engine，简称RBCC）和涡轮基组合动力循环发动机（turbine-based combined-cycle engine，简称TBCC）[1]。组合动力发动机本质上还是以未来飞行器宽速域和宽空域的需求为背景而产生的，图1是各类发动机适用的飞行速域和空域范围，可以看出组合动力发动机的优势是十分明显的。图1 各类发动机比冲随飞行马赫数的变化图2是RBCC发动机结构图 ，下面介绍一下火箭基组合动力循环发动机（RBCC）的工作过程：RBCC工作过程一般分为火箭引射模态（eject mode），亚燃冲压模态（ramjet mode），超燃冲压模态（scramjet mode），和纯火箭模态（pure rocket mode）四种模态。由文献[2]，四类模态的工作马赫数Ma范围分别是：火箭引射模态（eject mode）：Ma∈[0，3]亚燃冲压模态（ramjet mode）：Ma∈[3，7]超燃冲压模态（scramjet mode）：Ma∈[7，12]和纯火箭模态（pure rocket mode）：Ma>12可见RBCC可以从零起飞，首先起动火箭引射模态（0 西工大牵头研制的「飞天一号」发射成功，这一火箭取得了哪些技术突破？ - 知乎首页知乎知学堂发现等你来答切换模式登录/注册航天西北工业大学运载火箭航空航天西工大牵头研制的「飞天一号」发射成功，这一火箭取得了哪些技术突破？记者5日从西北工业大学获悉，由该校航天学院空天组合动力创新团队牵头研制的“飞天一号”火箭冲压组合动力4日在西北某基地发射成功。 [图片] 据介绍，这是…显示全部 关注者367被浏览455,423关注问题写回答邀请回答好问题 193 条评论分享47 个回答默认排序亨瑞瑞好好学习，天天向上 关注说“火箭”我觉得不准确，外媒说“导弹”更扯淡。我认为应该是“高超音速实验飞行器”这东西，军用就是高超音速武器——就是一个月前马督工引发“小”论战的那个玩意儿。引发睡前消息439期网络大讨论的 AGM-183A这东西，民用就是空天飞机——马斯克的“星舰“运载能力出众，眼瞅就上天了，我们目前还没有，不少网友认为，空天飞机就是我们可以弯道超车——或者至少作为追赶的“杀手锏”。“腾云工程”空天飞机妥妥的国之重器。报道的字不多——作为文傻，综合了一下。这次实验的“煤油燃料火箭冲压组合循环发动机“是什么？空天飞机不能只用传统的航空发动机——涡喷和涡扇发动机面对2.5马赫的来流速度，进气道和叶片间会产生很大的冲压效应。进气与压缩效率急剧下降——SR-71和米格25在3马赫的速度下发动机都采用了旁通路导气，叶片锁死或怠速，改用冲压模式工作；空天飞机也不能只用火箭发动机——本来发展空天飞机就是因为火箭的燃料效率太低了，在大气层内飞行，氧化剂就不用自己带了（有人说空天飞机比可回收火箭没有优势，也未尽然，即使成本没优势，在高超音速武器应用比弹道导弹强）所以理论上讲，空天飞机的发动机应该综合涡轮发动机——火箭发动机——冲压发动机。其中火箭发动机用于在涡轮发动机关闭后，冲压发动机启动前提供足够的来流速度，保证冲压发动机的启动。冲压发动机原理图具体实现方面，有以下技术路径：涡轮-冲压组合动力（TBCC）、火箭-冲压组合动力（RBCC）、涡轮-火箭组合动力（ATR）和三组合发动机（T/RBCC）。从报道描述看，这次西工大牵头研制的“飞天一号”实验的就是RBCC（Rocket-Based Combined Cycle，RBCC）。最核心的描述：冲压组合循环发动机火箭/亚燃、亚燃、超燃、火箭/超燃的多模态平稳过渡……核心是双模态超燃冲压发动机火箭/亚燃——从火箭加速后，在2-4马赫，启动亚燃冲压发动机，将高速气流降速增压，降至亚音速，燃烧后通过拉瓦尔喷管加速到超音速；亚燃——亚燃发动机稳定工作，继续加速；超燃——在来流速度达到5马赫以上时，亚燃模式转为超燃模式；火箭/超燃——这里我认为是以上整个过程，因为第一级是固体火箭，启动后不能停止，无法重复点火，而且图片来看，只有两级（百度说三级加液体，我没看出来第二级（及以上）在哪儿）。另外，使用煤油作为燃料是很重要的突破，其他答案有些了，我不再重复。（百度说31所提出了了固体火箭冲压基组合循环发动机(SRBCC：Solid-fuel Ram-Rocket Based Combined Cycle)新概念方案的设想和固液火箭冲压发动机(SRBLICC：Solid Rocket Based Liquid Injection CombinedCycle)——这次实验体的发动机准确描述还请大牛指正）后半句是（验证了）宽域综合能力，突破了热力喉道调节、超宽包线高效燃烧组织等关键技术热力喉道调节就是从亚燃到超燃两种模态变化的一个重要调节方式，通过热力喉道形成一个类似拉瓦尔喷管结构的热源，在亚燃模式下加速气流，在超燃模式下消失。由于人类对高超音速飞行还在探索中，相关数据、调节参数、特性的积累十分重要——必须依靠大量实验积累数据，完善模型。所以这里说关键技术突破毫不为过。事情说到这里就基本完了，但查资料的时候我发现了以下内容：箭体上印着的“北京动力机械研究所”是腾云工程的负责单位，在70年代就开始跟踪RBCC的重点技术90年代展开研究。而西北工业大学的相关团队，从一篇八股文的描述是“团队在1999年就敏锐的发现空天飞行器是未来国际航天大国竞争的新焦点，适用于空天飞行器的组合循环动力则是核心关键。这类发动机涉及到火箭发动机、涡轮发动机和冲压发动机的高度融合集成，难度极大。”也就是说，从跟踪到现在近50年，从研究到现在20多年！中国的航天人，数十年磨一剑！希望早日看到中国的空天飞机升空！参考资料及推荐阅读：世界瞭望塔：美国高超音速吸气式巡航导弹HAWC完成飞行测试超燃冲压发动机和普通的冲压发动机有什么区别（我国新型发动机已领先世界)全球首次！西工大冲压组合循环发动机：到底能干啥?究竟有多先进?【飞羽社】中国空天飞机发展计划——“腾云工程”到底是什么？_哔哩哔哩_bilibili编辑于 2022-07-07 09:48赞同 40052 条评论分享收藏喜欢收起永无赑屃 关注这个竟然没人关注？这个东西是真的牛逼。这次试验的飞天一号飞行器的飞行速度达到了高超音速，并且是采用的是吸气式的火箭冲压组合发动机。这意味着我国的高超音速巡航导弹及高超音速飞机的动力技术已经进入实用化阶段。这一技术水平在国际上是遥遥领先的。这次飞天一号使用的燃料是煤油，不像国外进行同类试验使用的液氢燃料。液氢燃料需要极低温度来保存，使用条件苛刻而且非常危险，易燃易爆，只能用在民用上没有军用价值。而煤油这是一种航空飞行普遍使用的常规燃料，常温常压下就可以安全保存。这就意味着我国高超音速巡航导弹已经具备了工程化研制的技术条件，或许用不了很久就会进入装备阶段。此次飞天一号使用的发动机是火箭冲压组合发动机，这已经不是国内第一次进行组合动力的试验了，国内目前至少已经过两次高超音速组合动力的飞行试验。从这次试验使用的动力来看是一种RBCC的组合动力。RBCC是一种火箭发动机为基础的组合动力，以火箭发动机为基础再组合冲压发动机。因为对于吸气式高超音速动力来说，单独一种发动机无法满足整个飞行速度域内的全部工况，必须组合不同种类的动力才能满足技术要求。从飞天一号的结构设计上来看，第一级是火箭发动机，也是助推器，第二级的飞行器才是本次试验的重点。第二级头部下方是有进气口的，也就是说第二级是要在大气层内飞行的。这样可以借助大气层屏蔽红外特征，有利于躲开敌方的弹道导弹预警卫星。这一点对于高超音速巡航导弹提高突防能力至关重要。令人吃惊的是第二级居然没有气动舵面，第二级的弹体尾部上方和下方有喷口设计，这就意味着第二级是靠矢量推力进行飞行控制的。一方面可以减少飞行阻力，一方面可以减少热防护的技术难度。因为气动舵面比较薄，不容易采用冷却技术措施，热防护的技术难度非常大。像我国的DF17高超音速导弹采用的就是空气舵作为飞行控制手段。DF17的飞行速度也是超过7马赫的，可能在7到10马赫之间。也就是说我国的热防护技术对付7到10马赫的飞行速度是完全没问题的，即使是空气舵这种比较薄，难以采取冷却技术措施的结构也没问题。这说明飞天一号的第二级的飞行速度非常快，有可能远远超过7马赫。因为对于使用碳氢燃料的超燃冲压发动机来说速度达到7马赫左右工作效率就开始下降了，超过8马赫就很困难了。虽然采用液氢燃料的超燃冲压发动机没有这一问题，但液氢燃料非常不实用，使用煤油等碳氢燃料才是最理想的。官方消息提到，此次飞行试验动力系统经历了火箭/亚燃、亚燃、超燃、火箭/超燃等多模态的切换。可以说极为复杂，这就是吸气式高超音速动力技术难度最高的地方。整个飞行试验的状态可能是这样的，首先由第一级火箭发动机点火起飞将第二级加速到两马赫以上。然后一二级分离，第二级发动机进入亚音速燃烧状态，在到达4马赫左右时进气道内的燃烧状态从亚音速转换为超音速燃烧状态。在2马赫到4马赫这个阶段，发动机推力比较小，称为推力陷阱。有可能需要火箭发动机介入，以帮助越过推力陷阱阶段。从4马赫到7马赫是超燃冲压工作状态，在7马赫左右从超燃冲压工作状态转换为火箭发动机工作状态。当使用火箭发动机工作时并不需要空气，此时是可以飞出大气层的。并且有可能达到10马赫甚至更高的飞行速度，这取决于燃料携带量和火箭发动机的工作时间。在整个飞行过程中要跨过多个工作模态。在不同的工作模态下，需要对热力喉道形状和截面进行调节，以满足不同模态的需求，实现在超宽包线内高效组织燃烧。飞天一号的动力不光可以用在高超音速巡航导弹上，也可以用在高超音速飞机上，以及两级入轨的空天飞机上，作为第一级的动力。这就意味着我国在2030年前实现高超音速组合动力的实际装备已经胜券在握了。一小时实现核打遍全球的时代不远了！这样的技术水平已经处于世界领先水平，已经把身后的国家远远抛到十万八千里外。7月4号发射的这个官媒报道了出来，而7月3号发射了一个更加先进的，官方没有报道。内容：过于先进，不便展示编辑于 2022-07-07 09:30赞同 695127 条评论分享收藏喜欢 如何看待西安航天动力研究所研制的某组合动力发动机首飞成功？ - 知乎首页知乎知学堂发现等你来答切换模式登录/注册工程学发动机航天航空航天如何看待西安航天动力研究所研制的某组合动力发动机首飞成功？西安航天动力研究所：某组合动力发动机首飞成功 “西安航天动力研究所” 9月22日消息：该所研制的某组合动力发动机首飞成功。 2020年金秋清晨，位于我…显示全部 关注者1,636被浏览1,809,403关注问题写回答邀请回答好问题 3505 条评论分享79 个回答默认排序Rocket Docter三流大学在读博士，为中华民族伟大复兴而读书 关注这是国内组合动力的首飞，是一个历史性时刻，值得庆祝。国际上两类主流的组合动力发动机就火箭基组合动力循环发动机（rocket-based combined-cycle engine，简称RBCC）和涡轮基组合动力循环发动机（turbine-based combined-cycle engine，简称TBCC）[1]。组合动力发动机本质上还是以未来飞行器宽速域和宽空域的需求为背景而产生的，图1是各类发动机适用的飞行速域和空域范围，可以看出组合动力发动机的优势是十分明显的。 图1 各类发动机比冲随飞行马赫数的变化图2是RBCC发动机结构图 ，下面介绍一下火箭基组合动力循环发动机（RBCC）的工作过程：RBCC工作过程一般分为火箭引射模态（eject mode），亚燃冲压模态（ramjet mode），超燃冲压模态（scramjet mode），和纯火箭模态（pure rocket mode）四种模态。由文献[2]，四类模态的工作马赫数Ma范围分别是：火箭引射模态（eject mode）：Ma∈[0，3]亚燃冲压模态（ramjet mode）：Ma∈[3，7]超燃冲压模态（scramjet mode）：Ma∈[7，12]和纯火箭模态（pure rocket mode）：Ma>12可见RBCC可以从零起飞，首先起动火箭引射模态（0 马赫数0至8范围RBCC发动机特性 期刊网首页 所有 标题 作者 关键词 摘要 Doi 栏目 作者地址 基金 中图分类号 高级检索 {{subColumn.name}} 高级检索 集群首页 {{newsColumn.name}} {{subColumn.name}} 马赫数0至8范围RBCC发动机特性 文章导航 >
空气动力学学报
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2022 >
40(1): 165-180 上一篇 下一篇 刘昊. 马赫数0至8范围RBCC发动机特性[J]. 空气动力学学报, 2022, 40(1): 165−180. doi: 10.7638/kqdlxxb-2021.0132 引用本文: 刘昊. 马赫数0至8范围RBCC发动机特性[J]. 空气动力学学报, 2022, 40(1): 165−180 . doi: 10.7638/kqdlxxb-2021.0132 LIU H. Characteristic analysis of RBCC engine at Mach number 0 to 8[J]. Acta Aerodynamica Sinica, 2022, 40(1): 165−180. doi: 10.7638/kqdlxxb-2021.0132 Citation: LIU H. Characteristic analysis of RBCC engine at Mach number 0 to 8[J]. Acta Aerodynamica Sinica, 2022, 40(1): 165−180 . doi: 10.7638/kqdlxxb-2021.0132 马赫数0至8范围RBCC发动机特性 刘昊 西安航天动力研究所，西安 710100 详细信息
作者简介: 刘昊*（1984-），男，陕西蓝田人，研究员，研究方向：组合推进技术. E-mail：ganeland@163.com 中图分类号:
V430
收稿日期:
2021-07-21 修回日期:
2021-09-11 录用日期:
2021-09-13 刊出日期:
2022-02-25 Characteristic analysis of RBCC engine at Mach number 0 to 8 Hao LIU Xi’an Aerospace Propulsion Institute, Xi’an 710100, China Received Date:
21 July 2021
Revised Date:
11 September 2021
Accepted Date:
13 September 2021
Publish Date:
25 February 2022
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摘要 摘要: 为获得飞行马赫数Ma0 = 0～8 RBCC发动机特性及结构调节规律，基于试验数据，建立了采用控制体法，考虑热完全气体效应、化学平衡流动效应、黏性损失及热损失等影响的发动机特性分析模型，并通过发动机自由射流试验获得的推力、比冲数据对所建立的发动机特性分析模型进行确认。完成二元中心火箭布局变结构模型RBCC发动机火箭引射模态、火箭冲压模态及冲压模态特性仿真，定量获得了飞行动压、马赫数、攻角、当量比、火箭流量等因素变化对发动机性能影响；并针对给定模拟飞行弹道，完成Ma0 = 0～8 RBCC发动机特性计算，给出了进气道收缩比、燃烧室扩张比、尾喷管扩张比、发动机总面积比随飞行马赫数及工作模态变化规律。研究表明：1）火箭引射模态，马赫数每增加1，推力、比冲增加约18.2%，火箭推力增益增加约15%；2）火箭冲压模态，火箭流量越大，火箭推力增益越小，且获得正的火箭推力增益范围越窄；3）Ma0 = 2模态转换点，发动机性能及结构参数均存在间断，确保推力及结构参数的连续调节、匹配应是模态转换规律制定的关注点；4）模拟弹道下，进气道收缩比、燃烧室扩张比、尾喷管扩张比、发动机总面积比在Ma0 = 0～8范围内分别变化6.17、4.26、30.38、8.94倍。 关键词:
高马赫数
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RBCC
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发动机特性
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分析模型
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模态转换
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变结构
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结构调节规律
Abstract: In order to obtain the characteristics and structural regulation law of the RBCC (Rocket-Based Combined-Cycle) engine in a flight Mach number range of Ma0 = 0～8, an engine characteristic analysis model is established based on experimental data, which adopts the control volume method and takes into account the effects of heat complete gas, chemical equilibrium flow, viscosity loss, heat loss, etc. The model is then verified using the thrust and specific impulse data obtained in a free-jet experiment. Characteristics of the RBCC engine of two-dimensional variable structures are analyzed for three operation modes, i.e. the injector operation mode, the ramjet operation mode and the rocket & ramjet mode. Effects of the flight dynamic pressure, Mach number, angle of attack, equivalent ratio and rocket flow rate on the performance of the model engine are quantitatively obtained. A characteristic analysis is carried out for the Ma0 = 0～8 RBCC engine along a simulated trajectory, and variations with the flight Mach number at different engine operation modes are given for the air inlet contraction ratio, combustor expansion ratio, nozzle expansion ratio, total area ratio. The results show that: 1) Under the ejector operation mode, for every increasement of one in the flight Mach number, the engine thrust and specific impulse are increased by about 18.2%, and the rocket thrust gain is increased by about 15%. 2) Under the ramjet operation mode, with the increase of the rocket flow rate, the rocket thrust gain decreases, and the larger the rocket flow rate is, the narrower the positive rocket thrust gain range is. 3) During the operation mode transformation at Ma0 = 2, the performance and structural parameters of the engine are both discontinoues, thus special attention needs to be put for the continuous adjustment of the thrust and structural parameters of the engine. 4) Under the given simulated trajectory, the inlet contraction ratio, the combustor expansion ratio, the nozzle expansion ratio and the engine total area ratio vary 6.17, 4.26, 30.38, 8.94 times in the range of Ma0 = 0～8,respectively. Key words:
high Mach number
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RBCC
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engine characteristics
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analysis model
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mode transformation
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variable structure
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structural regulation law
HTML全文 图 1
发动机构型及截面定义
Figure 1.
Schematic of the engine configuration and cross section definition
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全尺寸图片 幻灯片 图 2
进气道总压恢复系数计算与试验结果对比
Figure 2.
Comparison of total pressure recovery coefficients of the air inlet between simulation and test data
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全尺寸图片 幻灯片 图 3
进气道流量系数计算与试验结果对比
Figure 3.
Comparison of mass flux coefficients of the air inlet between simulation and test data
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全尺寸图片 幻灯片 图 4
引射系数计算与CFD结果对比
Figure 4.
Comparison of injection coefficients between simulation and CFD data
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全尺寸图片 幻灯片 图 5
燃烧效率计算与试验结果对比
Figure 5.
Comparison of combustion coefficients between simulation and test data
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燃烧室总压恢复系数计算与试验结果对比
Figure 6.
Comparison of total pressure recovery coefficients of the combustor between simulation and test data
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全尺寸图片 幻灯片 图 7
尾喷管总压恢复系数计算与CFD结果对比
Figure 7.
Comparison of total pressure recovery coefficients of the nozzle between simulation and CFD data
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Ma0 = 4工况计算与自由射流试验结果对比
Figure 8.
Comparison between simulation and test data at Ma0 = 4
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Ma0 = 6工况计算与自由射流试验结果对比
Figure 9.
Comparison between simulation and test data at Ma0 = 6
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火箭引射模态发动机速度-火箭节流特性
Figure 10.
Velocity and rocket throttle characteristic at the ejector operation mode
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发动机进出口速度随飞行马赫数变化
Figure 11.
Variation of velocities at the engine inlet and outlet with the flight Mack number
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引射系数随飞行马赫数变化
Figure 12.
Variation of the injection coefficient with the flight Mach number
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发动机进出口动量随飞行马赫数变化
Figure 13.
Variation of momentum at the engine inlet and outlet with the flight Mach number
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火箭引射模态火箭推力增益随飞行马赫数变化
Figure 14.
Variation of the thrust gain with the flight Mach number at the ejector operation mode
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冲压模态发动机速度-攻角特性
Figure 15.
Velocity and attack angle characteristics of the engine at the ramjet operation mode
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冲压模态发动机速度-节流特性
Figure 16.
Velocity and throttle characteristics of the engine at the ramjet operation mode
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冲压模态发动机速度-高度特性
Figure 17.
Velocity and altitude characteristics at the ramjet operation mode
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火箭冲压模态发动机速度-攻角特性
Figure 18.
Velocity and attack angle characteristics at therocket & ramjet operation mode
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火箭冲压模态发动机速度-节流特性
Figure 19.
Velocity and throttle characteristics at therocket & ramjet operation mode
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全尺寸图片 幻灯片 图 20
火箭冲压模态发动机速度-高度特性
Figure 20.
Velocity and altitude characteristics at therocket & ramjet operation mode
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全尺寸图片 幻灯片 图 21
火箭冲压模态发动机速度-火箭节流特性
Figure 21.
Velocity and rocket throttle characteristics at the rocket & ramjet operation mode
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全尺寸图片 幻灯片 图 22
火箭冲压模态火箭推力增益随飞行马赫数变化
Figure 22.
ΔFrocket&ramjet_mode - Ma0 relationship at the rocket & ramjet operation mode
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全尺寸图片 幻灯片 图 23
模拟弹道下发动机性能
Figure 23.
Engine performance along the simulated trajectory
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全尺寸图片 幻灯片 图 24
模拟弹道下进气道收缩比随飞行马赫数变化
Figure 24.
Variation of the air-inlet contraction ratio with the flight Mach number along the simulated trajectory
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全尺寸图片 幻灯片 图 27
模拟弹道下发动机总面积比随飞行马赫数变化
Figure 27.
Variation of the engine total area ratio with the flight Mach number along the simulated trajectory
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全尺寸图片 幻灯片 图 25
模拟弹道下燃烧室扩张比随飞行马赫数变化
Figure 25.
Variation of the combustor expansion ratio with the flight Mach number along the simulated trajectory
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全尺寸图片 幻灯片 图 26
模拟弹道下尾喷管扩张比随飞行马赫数变化
Figure 26.
Variation of the nozzel expansion ratio with the flight Mach number along the simulated trajectory
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全尺寸图片 幻灯片 表 1
火箭引射模态典型点推力增益
Table 1.
Thrust gain at typical conditions of the ejector operation mode
${\dot m_{{\rm{rocket}}} }$/(kg·s–1)MaminΔFejector_modeΔF/ΔMaMa0 = 0MaminMa0 = 2 100.3837.6%31.3%66.1%14.3%200.3630.1%24.6%59.7%14.8%300.3526.5%21.4%56.6%15.1%400.3424.1%19.2%54.4%15.2%500.3422.5%17.7%52.9%15.2%
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火箭冲压模态典型点推力增益
Table 2.
Thrust gain at typical conditionsof the rocket & ramjet mode
${\dot m_{{\rm{rocket}}} }$/(kg·s–1)MamaxΔFejector_modeMa0 = 2MamaxMa0 = 8 53.72.3%8.6%3.6%104.2–3.8%4.4%0.9%154.6–7.1%2.2%–0.5%204.9–9.3%0.7%–1.5%
下载: 导出CSV
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Figure 1. Schematic of the engine configuration and cross section definition
Figure 2. Comparison of total pressure recovery coefficients of the air inlet between simulation and test data
Figure 3. Comparison of mass flux coefficients of the air inlet between simulation and test data
Figure 4. Comparison of injection coefficients between simulation and CFD data
Figure 5. Comparison of combustion coefficients between simulation and test data
Figure 6. Comparison of total pressure recovery coefficients of the combustor between simulation and test data
Figure 7. Comparison of total pressure recovery coefficients of the nozzle between simulation and CFD data
Figure 8. Comparison between simulation and test data at Ma0 = 4
Figure 9. Comparison between simulation and test data at Ma0 = 6
Figure 10. Velocity and rocket throttle characteristic at the ejector operation mode
Figure 11. Variation of velocities at the engine inlet and outlet with the flight Mack number
Figure 12. Variation of the injection coefficient with the flight Mach number
Figure 13. Variation of momentum at the engine inlet and outlet with the flight Mach number
Figure 14. Variation of the thrust gain with the flight Mach number at the ejector operation mode
Figure 15. Velocity and attack angle characteristics of the engine at the ramjet operation mode
Figure 16. Velocity and throttle characteristics of the engine at the ramjet operation mode
Figure 17. Velocity and altitude characteristics at the ramjet operation mode
Figure 18. Velocity and attack angle characteristics at therocket & ramjet operation mode
Figure 19. Velocity and throttle characteristics at therocket & ramjet operation mode
Figure 20. Velocity and altitude characteristics at therocket & ramjet operation mode
Figure 21. Velocity and rocket throttle characteristics at the rocket & ramjet operation mode
Figure 22. ΔFrocket&ramjet_mode - Ma0 relationship at the rocket & ramjet operation mode
Figure 23. Engine performance along the simulated trajectory
Figure 24. Variation of the air-inlet contraction ratio with the flight Mach number along the simulated trajectory
Figure 27. Variation of the engine total area ratio with the flight Mach number along the simulated trajectory
Figure 25. Variation of the combustor expansion ratio with the flight Mach number along the simulated trajectory
Figure 26. Variation of the nozzel expansion ratio with the flight Mach number along the simulated trajectory
Figure FIG. 1347..
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支板式RBCC零速与非零速条件引射特性分析-《火箭推进》 航天推进技术研究院主办 smallTopNav 首 页期刊简介文章查询上期目录下期目录刊期索引下载排行作者园地投稿须知期刊订阅联系我们English 上一篇下一篇 PDF下载 分享 [1]万冰,陈军,白菡尘.支板式RBCC零速与非零速条件引射特性分析[J].火箭推进,2022,48(06):74-84. WAN Bing,CHEN Jun,BAI Hanchen.Analysis on the ejecting characterictics of a struct-based RBCC at zero speed and non-zero speed condition[J].Journal of Rocket Propulsion,2022,48(06):74-84. 点击复制 支板式RBCC零速与非零速条件引射特性分析 《火箭推进》[ISSN:1672-9374/CN:CN 61-1436/V]
卷: 48 期数: 2022年06期 页码: 74-84 栏目: 目次 出版日期: 2022-12-25 Title: Analysis on the ejecting characterictics of a struct-based RBCC at zero speed and non-zero speed condition 文章编号: 1672-9374(2022)06-0074-11 作者: 万冰; 陈军; 白菡尘 (中国空气动力研究与发展中心空天技术研究所高超声速冲压发动机技术重点实验室,四川绵阳621000) Author(s): WAN Bing; CHEN Jun; BAI Hanchen (Science and Technology on Scramjet Laboratory, Aerospace Technology Institute of CARDC, Mianyang 621000, China) 关键词: RBCC 引射 流场演化 速度条件 流量控制机制 Keywords: RBCC ejection process flow structure evolvement speed condition mass flow rate control mechanism 分类号: V235.21 文献标志码: A 摘要: 为获得来流速度条件对RBCC发动机引射特性的影响,采用数值模拟方法研究了支板式RBCC构型在亚声速范围的引射流场特性。结果表明:有/无速度条件的引射过程均存在两个阶段,即反压影响阶段和自维持阶段,在自维持阶段,内流道存在全为超声速的截面,环境压力不会对二次流流量产生影响 当处于反压影响阶段,由于环境压力不同,不同速度条件的内流场存在差异,二次流流量也存在差别,随马赫数增加,RBCC所引射的空气流量增加 当处于自维持阶段时,不同速度条件的内流场十分相似,二次流流量也基本相同,说明二次流总温、总压相同时,马赫数对引射过程没有影响,有速度条件的引射过程可以等效为相同总温、总压的零速引射过程,这为有速度条件的二次流流量评估以及试验来流参数配置提供了便利。 Abstract: In order to obtain the influence of the speed condition on the ejecting process of a RBCC(rocket based combined cycle)engine, numerical simulation was performed to investigate the ejector mode of a strut-based RBCC model at subsonic speed.The results show that the operation regimes contains two stages under the investigated speed conditions are back pressure dominating stage and self-sustaining stage.At self-sustaining stage, due to the existence of all-supersonic flow, the ambient pressure can not influence the ejecting process.When working at back pressure dominating stage, due to the different ambient pressure(the total pressure is the same), the internal flow pattern and the mass flow rate of the secondary flow are different at different Mach number while instead, the internal flow pattern and the mass flow rate of the secondary flow are almost the same at self-sustaining stage, which further shows the parameters of the secondary flow that determine the ejection process are the total pressure and temperature, irrelevant to Mach number.So, the ejection process at some speed condition can be equivalent to the ejection process at zero speed provided that their total pressure and temperature are the same, which provides the convenience for the assessment of the mass flow rate of the secondary flow at subsonic speed and the preparation for flow condition of wind tunnel test.
参考文献/References: [1] 龚春林,陈兵.组合循环动力在水平起降天地往返飞行器上的应用[J].科技导报,2020,38(12):25-32.[2] 张升升,郑雄,吕雅,等.国外组合循环动力技术研究进展[J].科技导报,2020,38(12):33-53.[3] 张玫,张蒙正,刘昊.火箭基组合循环动力研究进展[J].科技导报,2020,38(12):54-68.[4] 王亚军,何国强,秦飞,等.火箭冲压组合动力研究进展[J].宇航学报,2019,40(10):1125-1133.[5] BULMAN M,SIEBENHAAR A.Combined cycle propulsion:Aerojet innovations for practical hypersonic vehicles[C]//17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference.Reston,Virginia:AIAA,2011.[6] 张蒙正,张玫,严俊峰,等.RBCC动力系统工作模态问题[J].火箭推进,2015,41(2):1-6.ZHANG M Z,ZHANG M,YAN J F,et al.Discussion about work modal of RBCC power system[J].Journal of Rocket Propulsion,2015,41(2):1-6.[7] 白菡尘,陈军.双模态冲压发动机等效热力过程与性能关系原理[M].北京:国防工业出版社,2018.[8] SIEBENHAAR A,BONAR D,SARMONT E.The role of the strutjet engine in new global and space markets[EB/OL].https://www.semanticscholar.org/paper/The-Role-of-the-Strutjet-Engine-in-New-Global-and-Siebenhaar-Bonar/75c5 7488e9bd124296bc9a3565e0d0df69e00aca,1998.[9] QUINN J.ISTAR:Project status and ground test engine design[C]//39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit.Reston,Virginia:AIAA,2003.[10] TAKEGOSHI M,TOMIOKA S,UEDA S,et al.Performances of a rocket chamber for the combined-cycle engine at various conditions[C]//14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference.Reston,Virginia:AIAA,2006.[11] 黄生洪,TOHRU M,EIJIRO K.壅塞机制对 RBCC 引射模态推力增强性能的影响[C]//中国力学学会学术大会论文集.北京:中国力学学会,2005.[12] DIJKSTRA F,MAREE A,CAPORICCI M,et al.Experimental investigation of the thrust enhancement potential of ejector rockets[C]//33rd Joint Propulsion Conference and Exhibit.Reston,Virginia:AIAA,1997.[13] 阿勃拉莫维奇.实用气体动力学[M].梁秀彦,译.北京:高等教育出版社,1955.[14] FABRI J,SIESTRUNCK R.Supersonic air ejectors[J].Advances in Applied Mechanics,1958,5:1-34.[15] FABRI J,PAULON J.Theory and experiments on supersonic air-to-air ejectors[R].NACA T-M1410.[16] 朱也夫В С,马卡伦В С.冲压和火箭冲压发动机原理[M].北京:国防工业出版社,1975.[17] AOKI S,LEE J,MASUYA G,et al.Aerodynamic experiment on an ejector-jet[J].Journal of Propulsion and Power,2005,21(3):496-503.[18] TANI K,KANDA T,TOKUTOME S.Aerodynamic characteristics of the combined cycle engine in an ejector jet mode[C]//43rd AIAA Aerospace Sciences Meeting and Exhibit.Reston,Virginia:AIAA,2005.[19] 刘洋,何国强,刘佩进,等.一次火箭流量对RBCC性能影响的数值和实验研究[J].固体火箭技术,2008,31(5):439-444.[20] 李强,刘佩进,李江,等.来流马赫数对引射火箭引射量的影响研究[J].固体火箭技术,2009,32(3):254-257.[21] 刘晓伟,石磊,刘佩进,等.RBCC发动机引射模态进气道特性研究[J].固体火箭技术,2016,39(5):601-605.[22] SHI L,LIU X W,HE G Q,et al.Numerical analysis of flow features and operation characteristics of RBCC inlet in ejector mode[J].Acta Astronautica,2016(127):182-196.[23] LEHMAN M,PAL S,SANTORO R.Experimental investigation of the RBCC rocket-ejector mode[C]//36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit.Reston,Virginia:AIAA,2000.[24] 陈军,白菡尘.稳态引射过程自维持临界截面研究[J].推进技术,2022,43(5):74-83.
备注/Memo 收稿日期:2022-03-05 修回日期:2022-04-05基金项目:国家重点实验室基金(STS/MY-ZY-2020-004)作者简介:万冰(1994―),男,博士研究生,研究领域为组合循环发动机设计理论。通信作者:白菡尘(1965―),女,博士,研究员,研究领域为吸气式发动机内流空气动力学。
更新日期/Last Update:
1900-01-01 版权所有© 2011 《火箭推进》编辑部 ICP备00000000号Tel: 029-85208348 Email:hjtj@casc11.comPowered By 西安三才科技实业有限公司 029-88222992 www.samsoncn.com 组合循环推进系统燃料消耗模型及优化分析 引用本文
计自飞,
王兵,
张会强.
组合循环推进系统燃料消耗模型及优化分析[J]. 清华大学学报(自然科学版), 2017, 57(5): 516-520.
JI Zifei,
WANG Bing,
ZHANG Huiqiang.
Optimal analysis of the fuel consumption of combined cycle propulsion systems[J]. Journal of Tsinghua University (Science and Technology), 2017, 57(5): 516-520.
组合循环推进系统燃料消耗模型及优化分析
计自飞 , 王兵 , 张会强
清华大学 航天航空学院, 北京 100084
收稿日期：2015-12-11
基金项目：清华大学自主科研计划（20141081217）
作者简介：计自飞(1991—), 男, 博士研究生
通信作者：王兵, 副教授, E-mail:wbing@tsinghua.edu.cn 摘要：根据飞行任务要求，准确计算出飞行器所需的燃料消耗是推进系统设计的前提。该文针对火箭基组合循环动力（RBCC）推进方式，并以“地面起飞—巡航—滑翔着陆”的高超音速飞行器为研究对象，采用理论分析的方法建立了燃料消耗的计算模型，并提出一种以最小燃料消耗为目标的多参数优化方法。周期跳跃式巡航飞行器燃料消耗的研究结果表明：随着巡航初速度、爬升段航迹倾角、巡航轨迹角的增加，燃料消耗量增加；随着飞行动压的增加，燃料消耗量先减小后阶跃式增加。优化分析结果表明：对于起飞质量100 t、2 h全球到达的RBCC组合动力高超音速飞行器，在升阻比为4时巡航跳跃周期数为46，最小燃料消耗量约为32 t。研究结果表明该燃料消耗分析方法合理、可行，为高超音速飞行器及组合循环动力推进系统的工程设计提供了依据。
关键词：组合循环推进 高超音速飞行 燃料消耗 优化方法
Optimal analysis of the fuel consumption of combined cycle propulsion systems
JI Zifei,
WANG Bing,
ZHANG Huiqiang
School of Aerospace Engineering, Tsinghua University, Beijing 100084, China Abstract: The propulsion system fuel consumption must be accurately predicted for aircraft missions. A theoretical analysis of a hypersonic aircraft with a "boost-cruise-glide" flight mission profile powered by a rocket based combined cycle (RBCC) engine is used to predict the fuel consumption of the aircraft and to optimize the fuel consumption. The fuel consumption analysis of periodic hypersonic cruise trajectories shows that the fuel consumption decreases with increasing initial cruise velocity, larger flight-path angles and larger flight-path angles for cruising. As the flight dynamic pressure increases, the fuel consumption first decreases but then increases with a step change. The optimization results show that the hypersonic cruise stage should have 46 skip-periods with a minimum fuel consumption of about 32 tons for a two-hour global-reaching hypersonic aircraft with an initial weight of 100 tons and a lift-drag ratio of 4. The optimal results show that the fuel consumption prediction model is reasonable. The present study can guide the design of combined cycle propulsion systems.
Key words:
combined cycle propulsion
hypersonic flight
fuel consumption
optimization method
飞行Mach数Ma在5~6以上的高超音速飞行器具有优越的高空高速特性、灵活的机动性、高的突防概率以及可重复利用从而成本低等优点，因而高超音速飞行器受到各军事强国的重视。高超音速飞行器具有工作包线宽、飞行工况复杂多变等特点，目前的单一类型动力与推进系统难以满足其飞行要求。为此，人们提出了组合循环动力系统，将吸气式航空涡轮发动机、冲压发动机和自带推进剂的火箭发动机等不同类型的推进方式组合起来，实现多个不同的工作模态，以满足高超音速飞行器在不同工作阶段对推进系统的要求[1-2]。 由于基础发动机的类型不同，组合循环动力的类型也不同，通常有火箭基组合循环(rocket based combined cycle, RBCC)发动机和涡轮基组合循环(turbine based combined cycle, TBCC)发动机等。针对RBCC，国内外学者开展了大量的理论和实验研究[3-6]，并取得了一系列成果，为组合循环动力方式的工程实践和关键技术突破打下了坚实的基础。图 1给出了一种RBCC发动机结构示意图，它由冲压发动机和安装于其中心锥体中的火箭发动机组成。液体火箭发动机在工作时产生高温高压燃气并与引射的空气进行混合。目前来看，RBCC组合动力系统的工作模态包括引射火箭模态(Ma < 3)、火箭-亚燃冲压、亚燃冲压(Ma在3~6)、火箭-超燃冲压、超燃冲压、火箭模态(Ma>10) 等多个工作模态[3, 5]。上述各模态之间转换点设计的原则是在满足推力需求前提下，比冲尽可能高。 图 1 火箭置于中心锥体的一种RBCC结构示意图
图选项 燃油经济性是高超音速飞行器的一项重要评价指标，根据高超音速飞行任务的要求，准确分析并确定飞行器的燃料消耗是飞行器及其推进系统设计的前提条件。文[7]给出了稳态巡航条件下高超音速飞行器最省燃料的飞行参数；文[8]提出了稳态巡航和周期巡航两种情形燃料消耗的计算方法。然而，针对高超音速飞行器“地面起飞—巡航—滑翔着陆”飞行全过程的燃料消耗计算方法和优化分析鲜见报道。燃料消耗与飞行参数密切相关，确定最小的燃料消耗属于有约束的最小值优化问题。然而，由于目标函数复杂，影响参数多，难以直接利用现有的遗传算法、粒子群算法等优化方法[9]获得优化分析结果。 本文针对具有4种工作模态的RBCC动力系统，建立了高超音速飞行过程的组合推进燃油消耗模型，并提出了一种以最小化燃料消耗为目标基于逐次逼近全局搜索的多参数优化方法。应用该模型和优化方法，研究了主要参数对组合循环动力高超音速飞行过程及燃料消耗的影响。 1 RBCC发动机燃油消耗计算模型 参照文[10-13]的思路，本研究首先对RBCC发动机整体性能进行参数化分析，获得了比推力Fm和比冲Isp性能随飞行Ma数的变化，并与文[14-15]中的结果进行对比，如图 2所示。可见，本文结果与文献中的结果吻合很好。 图 2 RBCC发动机比推力以及比冲随飞行Ma数的变化
图选项 1.1 爬升段燃料消耗 对于按照等动压(动压q数值一般在37.3~50.0 kPa之间变化[1])飞行的高超音速飞行器，假定航迹倾角α和迎角β的变化规律已知，则飞行速度与飞行时间之间的关系是q、h(高度)、α、β的函数，v(t)=f(q, h, α, β)。结合爬升段结束时的飞行高度，可得爬升段总时间tboost。 飞行器所受的气动力由文[16-17]模型求得，根据飞行方程，可以推导推力随时间的变化情况F(t)。按照组合循环发动机的比冲与飞行Ma数的关系，用a表示加速度，则可得到爬升段的燃料消耗：
$ m\left( t \right) = {m_0}-\int_0^t {{{\dot m}_{{\rm{fuel}}}}} {\rm{d}}t, $
(1)
$ {m_{{\rm{fuel, boost}}}} = \int_0^{{t_{{\rm{boost}}}}} {\frac{F}{{g{I_{{\rm{sp}}}}}}} {\rm{d}}t = \int_0^{{t_{{\rm{boost}}}}} {\frac{{D + m(g{\rm{sin}}\alpha + a)}}{{g{I_{{\rm{sp}}}}{\rm{cos}}\beta }}{\rm{d}}t} . $
(2) 其中：D为阻力，m为飞行器的瞬时质量，m0为飞行器的起飞质量，${\dot m_{{\rm{fuel}}}} $为单位时间消耗的燃料质量，mfuel, boost表示爬升段消耗的燃料质量，g为重力加速度。 1.2 巡航段燃料消耗 巡航态是高超音速飞行器的主要工作状态，通常有两种形式：稳态平飞巡航和周期跳跃巡航。对于稳态巡航，可认为其飞行高度和速度均为定值，航迹倾角α=0。只要给定迎角β，则气动参数可确定，稳定巡航状态可解，在该状态下的燃料消耗可求，
$ \begin{array}{l} {m_{{\rm{fuel, cruise}}}} = \int_0^{{t_{{\rm{cruise}}}}} {\frac{F}{{g{I_{{\rm{sp}}}}}}} {\rm{d}}t = \frac{{D\left( \beta \right)}}{{g{I_{{\rm{sp}}}}{\rm{cos}}\beta }}{t_{{\rm{cruise}}}} = \\ \frac{{D\left( \beta \right)}}{{g{I_{{\rm{sp}}}}{\rm{cos}}\beta }}\frac{{{R_{{\rm{cruise}}}}}}{{Ma \cdot c}}\left( {1 + \frac{h}{{{r_{\rm{E}}}}}} \right). \end{array} $
(3) 其中：tcruise代表稳态巡航时间，Rcruise代表稳态巡航航程，rE为地球半径，c为当地音速。 周期跳跃巡航指的是爬升段结束后，飞行器开始无动力巡航，直到高度降到某一设定值发动机重新点火，当飞行器加速到巡航初速后，发动机再次关闭，如此周而复始。对于周期跳跃式巡航，在计算燃料消耗时，根据在每个周期内已知的飞行器初速度、末速度、始末工作高度，得出发动机工作时间${\bar t_i} $，以及在此期间飞行器速度变化v(t)，随后由迎角β确定气动力，最后由动力学方程得到推力F(t)，则燃料消耗可求，
$ {m_{{\rm{fuel, cruise}}}} = \sum\limits_i {{m_i} = } \sum\limits_i {\int_0^{{{\bar t}_i}} {\frac{F}{{g{I_{{\rm{sp}}}}}}} } {\rm{d}}t. $
(4) 1.3 着陆段燃料消耗 高超音速飞行器采用滑翔方式着陆，并且该阶段无燃料消耗。对于着陆段，人们关注的重点在于如何在给定的高度和初速度条件下获得最远的航程[18]，同时需要考虑的是高超音速飞行器需要在规定时间内完成预定任务。 2 针对组合循环推进及跳跃式巡航方式的最小燃油消耗优化方法 2.1 优化方法 基于最小燃料消耗的参数优化问题属于有约束条件的非线性最小值问题，为了计算方便，同时考虑到约束条件相对简单，可通过将约束条件加入目标函数中而使问题转化为无约束非线性最小值问题。针对此类问题的现有算法，往往只能找到局部最优值。为了能够简单快速地得到优化问题的全局最优值，本文提出一种逐次逼近全局搜索的优化算法，该算法的基本流程如图 3所示。 图 3 优化方法流程图
图选项 2.2 优化案例分析 针对某2 h全球到达的RBCC高速飞行器，以燃料消耗最小为目标进行优化计算。要求飞行器从地面起飞，起飞质量100 t，航程不短于2×104km，任务时间不超过2 h，采用跳跃式高超音速巡航轨迹。考虑到方案的可行性，对一些参数进行必要的限制，如表 1所示。涡轮发动机燃料为航空煤油，冲压发动机燃料为液氢，火箭发动机推进剂为液氢/液氧。 表 1 优化设计参数的限制条件 参数 范围 升阻比L/D 2~4 飞行器Ma < 15 冲压工作高度 < 40 km 表选项 爬升段按照等动压飞行，假定航迹倾角α=const，首先得到F(t)和v(t)，从而得到迎角。对于周期跳跃式巡航的运动学分析参考文[19]，将每个巡航周期分为弹道段和跳跃段两部分。考虑到发动机每次开机时间非常短，因此忽略加速时间。 图 4和5分别给出了不同升阻比条件下巡航初速度v0和巡航轨迹角θ对燃料消耗量mfuel、巡航周期数N和总任务时间t的影响。燃料消耗量随着巡航初速度的增加而增加，总飞行时间和巡航周期数随着巡航初速度的增加而减小。由于巡航初速度增加，爬升末速度和滑翔初速度增加，使得这两段航程增加，巡航周期数便随之减小。 图 4 燃料消耗、巡航周期数和总时间随巡航初速度的变化
图选项 图 5 燃料消耗、巡航周期数和总时间随巡航轨迹角的变化
图选项 燃料消耗量随着巡航轨迹角的增加呈阶梯状增加，巡航周期数和总飞行时间随巡航轨迹角的增加呈阶梯状减小。任务时间、燃料消耗曲线与巡航周期数曲线在相同的巡航轨迹角度值处出现阶跃，可见燃料消耗和任务时间出现阶跃的原因是巡航周期数的变化。 爬升段的飞行动压以及航迹倾角对燃料消耗、巡航周期数和总飞行时间也有显著影响。总之，飞行器的升阻比越大，燃料消耗量越少。巡航周期数对巡航轨迹角的变化更加敏感，而对爬升段航迹倾角的变化则不敏感。 3 高超音速飞行器最小燃料消耗的多参数优化分析 3.1 多参数优化结果与分析 利用第2节的优化算法进行多参数优化。依据图 4和5的数据，确定首次搜索的步长，该步长应当小于参数引起巡航周期数变化的量，否则优化结果将陷入局部最优。表 2给出了一组以RBCC为动力的飞行参数优化结果。 表 2 以最省燃料为目标优化飞行参数 状态参数 优化结果 总燃料消耗量mfuel/t 31.66 爬升段航迹倾角α/(°) 5 巡航轨迹角θ/(°) 11 巡航初速度v0 /(m·s-1) 3 780 飞行动压q0/kPa 38.0 巡航周期数 46 升阻比L/D 4 表选项 在表 2给定的飞行状态下，选用RBCC作为组合动力高超音速飞行器的飞行过程如图 6所示。 图 6 RBCC动力飞行器飞行过程优化结果示意图
图选项 优化后，飞行器的飞行过程如下：以5°的航迹倾角从地面起飞，采用等动压飞行方式迅速爬升，经过73.75 s飞行器达到40 km高空，航程达到458.95 km，消耗燃料31.18 t。爬升段结束开始周期巡航，每个巡航周期包括弹道段和跳跃段两个阶段，在弹道段飞行器以惯性继续爬升一定高度，随后开始下降，当高度再次回到40 km时，弹道段结束，同时跳跃段开始，当飞行器继续下降一定高度后，发动机再次点火工作，飞行器开始上升，同时巡航速度回到弹道段初速度，至此跳跃段结束。每个弹道段用时56.08 s，航程208.08 km；每个跳跃段用时76.67 s，航程142.23 km。巡航段经过46个周期(其中弹道段47个)，用时6.16×103 s, 航程1.63×104 km, 消耗燃料0.48 t。巡航结束，飞行器开始滑翔着陆，滑翔段用时903.25 s，航程3.31×103km。整个飞行过程用时7.14×103s，总航程R达到2.01×104km，共消耗燃料31.66 t，如图 6所示。 3.2 不同推进方式的比较分析 分别选择火箭发动机、火箭基组合发动机和涡轮基组合发动机进行优化设计，设定相同的飞行轨迹，最小燃料消耗量如图 7所示。 图 7 3种推进方式的最小燃料消耗量对比
图选项 由图 7可见，采用火箭发动机时燃料消耗量最大，组合循环动力在节省燃料方面具有明显优势，尤其是涡轮基组合发动机，但其目前技术成熟度较低。火箭基组合发动机在引射模态的燃料消耗较大，这是因为火箭发动机的比冲远小于涡轮发动机，而在其他工作模态，火箭基与涡轮基组合发动机的燃料消耗量基本相当。因此，对于RBCC而言，引射火箭是当前值得深入研究的关键技术。 4 结论 本文以采用RBCC组合循环发动机为动力并基于“地面起飞—巡航—滑翔着陆”飞行轨迹的高超音速飞行器为研究对象，建立了不同飞行阶段燃料消耗的数学模型，并提出了一种以最小燃料消耗量为优化目标的多参数逐次逼近全局搜索优化方法。以全球2 h到达的跳跃式高超音速巡航飞行器的燃料消耗为例，进行了多参数优化分析。 研究表明，本文的燃料消耗计算方法和优化算法合理、可行，适用于周期跳跃巡航和平飞稳态巡航两类飞行方式。对于起飞质量100 t、采用跳跃式巡航轨迹、全球2 h到达的组合动力高速飞行器，最小的燃料消耗量约为32 t，跳跃周期数为46。相比于火箭发动机，组合循环动力发动机在节省燃料方面优势显著；在相同飞行条件下，火箭基组合循环引射模态的燃料消耗远超过涡轮基组合循环动力引射模态燃料消耗，使得其总体燃料消耗量较大。 参考文献 [1]
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doi:10.16511/j.cnki.qhdxxb.2017.22.030 0 文章信息
计自飞,
王兵,
张会强
JI Zifei,
WANG Bing,
ZHANG Huiqiang
组合循环推进系统燃料消耗模型及优化分析 Optimal analysis of the fuel consumption of combined cycle propulsion systems 清华大学学报(自然科学版), 2017, 57(5): 516-520. Journal of Tsinghua University (Science and Technology), 2017, 57(5): 516-520. 收稿日期: 2015-12-11 工作空间 火箭基组合循环动力研究进展 加入收藏 | 设为首页 首页 期刊简介 编委会 大事记 投稿指南 期刊订阅 下载中心 项目合作 广告合作 联系我们 English 科技导报 2020, Vol. 38 Issue (12): 54-68 DOI: 10.3981/j.issn.1000-7857.2020.12.005
专题：先进组合动力飞行器 本期目录 | 过刊浏览 | 高级检索
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火箭基组合循环动力研究进展 张玫, 张蒙正, 刘昊 西安航天动力研究所, 西安 710100 Progress and analysis of rocket based combined cycle(RBCC) propulsion system ZHANG Mei, ZHANG Mengzheng, LIU Hao Xi'an Aerospace Propulsion Institute, Xi'an 710100, China 摘要 图/表 参考文献 相关文章
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摘要 重复使用是运载器发展的必然途径，火箭基组合循环发动机是可重复运载器动力的重要方向。介绍了国内外典型RBCC发动机方案和关键技术研究现状，分析了RBCC发动机的主要技术特点及应用前景。可以弥补火箭或冲压单一类型发动机功能或者性能的不足，具有火箭大推重比、冲压高比冲的特点，是RBCC动力系统区别于其他发动机的重要特征。结合当前技术水平，灵活运用组合发动机的特点，形成不同的发动机方案，适用不同运载任务要求，是RBCC动力系统研究的重要思路。中国应加快RBCC发动机应用论证和关键技术攻关，形成技术方案，为可重复使用运载器长远发展做出贡献。 服务 把本文推荐给朋友 加入我的书架 加入引用管理器 E-mail Alert RSS 作者相关文章 张玫 张蒙正 刘昊 关键词 ：
火箭基组合循环发动机,
可重复使用运载器,
RBCC动力
Abstract：Reuse is a promising technology for launch vehicles. Rocket based combined cycle(RBCC) propulsion system is a good choice for reusable launch vehicles. In this article the development status of RBCC engine and key technological issues are described, possible application areas, and development path are analyzed as well. To make full use of the advantages of ramjet and rocket engine and remedy function disfigurement and performance deficiency of LRE(改为全称) or ramjet distinguishes RBCC engine from other power engines. On the basis of current technologies, the RBCC engine solution formed by different working modals and engine configurations is the developmental route of RBCC power system to meet the needs of fly mission. Key words：
rocket based combined cycle(RBCC) propulsion system
launch vehicle
RBCC power 收稿日期: 2019-11-09
作者简介: 张玫,高级工程师,研究方向为液体火箭发动机,电子信箱:1308328295@qq.com 引用本文:
张玫, 张蒙正, 刘昊. 火箭基组合循环动力研究进展[J]. 科技导报, 2020, 38(12): 54-68.
ZHANG Mei, ZHANG Mengzheng, LIU Hao. Progress and analysis of rocket based combined cycle(RBCC) propulsion system. Science & Technology Review, 2020, 38(12): 54-68.
链接本文: http://www.kjdb.org/CN/10.3981/j.issn.1000-7857.2020.12.005 或 http://www.kjdb.org/CN/Y2020/V38/I12/54 [1] 张蒙正, 李平, 陈祖奎. 组合循环动力系统面临的挑战及前景[J]. 火箭推进, 2009, 35(1):1-8.
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电话：010-62138113，传真：010-62138113 电子信箱：kjdbbjb@cast.org.cn Thrust performance of the rocket-based combined-cycle engine under ejector mode | AIP Advances | AIP Publishing Skip to Main Content Close Publishers AIP Publishing Physics Today Acoustical Society of America American Association of Physics Teachers American Crystallographic Association, Inc. AVS: Science and Technology of Materials, Interfaces and Processing Chinese Physical Society Hefei General Machinery Research Institute Laser Institute of America The Society of Rheology Tianjin University Publications Journals Physics Today Conference Proceedings Books Special Topics Authors Librarians About Search Dropdown Menu header search search input Search input auto suggest filter your search All ContentAIP Publishing PortfolioAIP Advances Search Advanced Search |Citation Search User Tools Dropdown
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August 2023 Previous Article Next Article NOMENCLATURE Abbreviations Variables Subscripts I. INTRODUCTION II. THE DEFINITION OF CHOKING IN THE RBCC EJECTOR MODE III. IDEAL THERMODYNAMIC CYCLE ANALYSIS OF THE RBCC WITH THE ENGINE EJECTOR MODE A. Assumptions of the thermodynamic cycle analysis B. The thermodynamic cycle process analysis C. The energy conversion analysis and the influence factors IV. MATHEMATICAL MODELING FOR THE RBCC EJECTOR MODE A. Assumptions used in the quasi-one-dimensional model B. The definition of some important parameters C. Ejecting and mixing processes D. Flow choking state E. Process in the diffuser F. Secondary combustion process G. Expansion process in the nozzle V. RESULTS AND DISCUSSION A. Low Mach and low altitude flight conditions B. Low Mach and high altitude flight conditions C. High Mach and low altitude flight conditions D. High Mach and high altitude flight conditions E. Performance trend of the RBCC engine VI. CONCLUSION ACKNOWLEDGMENTS AUTHOR DECLARATIONS Conflict of Interest Author Contributions DATA AVAILABILITY REFERENCES Article Navigation Research Article| August 30 2023
Thrust performance of the rocket-based combined-cycle engine under ejector mode
Yizhi Yao (姚轶智)
0000-0001-5309-8361
;
Yizhi Yao (姚轶智)
(Conceptualization, Data curation, Formal analysis, Funding acquisition, Visualization, Writing – original draft, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Rui Gu (顾瑞)
0009-0009-5709-3061
;
Rui Gu (顾瑞)
(Data curation, Software, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Mingbo Sun (孙明波)
0000-0003-1676-4008
;
Mingbo Sun (孙明波)
a)
(Project administration, Resources, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China a)Authors to whom correspondence should be addressed: sunmingbo@nudt.edu.cn and lipeibo09@nudt.edu.cn
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This Site PubMed Google Scholar Peibo Li (李佩波)
0000-0003-3639-8551
;
Peibo Li (李佩波)
a)
(Funding acquisition, Project administration, Resources, Validation)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China a)Authors to whom correspondence should be addressed: sunmingbo@nudt.edu.cn and lipeibo09@nudt.edu.cn
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This Site PubMed Google Scholar Yuhui Huang (黄玉辉);
Yuhui Huang (黄玉辉)
(Funding acquisition, Resources, Writing – review & editing)
2Center for Project Management of Equipment Development Department, Beijing 110000, China
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This Site PubMed Google Scholar Bin An (安彬)
0000-0002-7511-3654
;
Bin An (安彬)
(Data curation, Formal analysis, Funding acquisition, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Jiaoru Wang (王教儒)
0000-0001-5346-0383
;
Jiaoru Wang (王教儒)
(Formal analysis, Writing – original draft, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Menglei Li (李梦磊)
0009-0006-9347-058X
;
Menglei Li (李梦磊)
(Visualization, Writing – original draft, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Taiyu Wang (王泰宇)
0000-0001-8236-2926
;
Taiyu Wang (王泰宇)
(Writing – original draft, Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Jikai Chen (陈纪凯)
0000-0003-4219-5398
Jikai Chen (陈纪凯)
(Writing – review & editing)
1Science and Technology on Scramjet Laboratory, College of Aerospace Science and Engineering, National University of Defense Technology, Changsha, Hunan 410073, China
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This Site PubMed Google Scholar Author & Article Information
a)Authors to whom correspondence should be addressed: sunmingbo@nudt.edu.cn and lipeibo09@nudt.edu.cn AIP Advances 13, 085036 (2023) https://doi.org/10.1063/5.0145047 Article history Received: February 02 2023 Accepted: August 07 2023 Split-Screen Views Icon Views Article contents Figures & tables Video Audio Supplementary Data Peer Review Open the PDF for in another window Share Icon Share Twitter Facebook Reddit LinkedIn Tools Icon Tools Reprints and Permissions Cite Icon Cite Search Site Citation Yizhi Yao, Rui Gu, Mingbo Sun, Peibo Li, Yuhui Huang, Bin An, Jiaoru Wang, Menglei Li, Taiyu Wang, Jikai Chen; Thrust performance of the rocket-based combined-cycle engine under ejector mode. AIP Advances 1 August 2023; 13 (8): 085036. https://doi.org/10.1063/5.0145047 Download citation file: Ris (Zotero) Reference Manager EasyBib Bookends Mendeley Papers EndNote RefWorks BibTex toolbar search Search Dropdown Menu toolbar search search input Search input auto suggest filter your search All ContentAIP Publishing PortfolioAIP Advances Search Advanced Search |Citation Search The Rocket-Based Combined-Cycle (RBCC) engine integrates the preponderance of ramjet and rocket engines. It can perform excellently at a lower Mach number than a ramjet while consuming less fuel than the rocket. The higher specific impulse under lower flight Mach conditions guarantees the competitiveness and application prospect of an RBCC engine for reusable space transportation and hypersonic cruise vehicles. With a wide range of Mach numbers, the flow choking between primary and secondary streams in the inner flow passage of the engine becomes complicated. The flow choking not only affects the mass flow ratio between the air stream and rocket plume but also determines the thrust performance of the engine. However, the relationship between flow-choking states and thrust performance has not been revealed. This investigation aimed to provide a better design reference for the RBCC trajectory design and a basis for the RBCC engine geometry design so that the thrust performance under different flow choking states was studied. The findings indicate that the engine is more favorable in the supersonic regime status than others. At the premise of the lower total pressure air stream condition, the thrust of the RBCC engine is attributed to the rocket pressurization effect. On the contrary, its thrust is ascribed to the air stream when the air stream total pressure is higher. Besides, the thrust augmentation of the RBCC engine is essentially due to the larger entrainment ratio. There is a rocket mass flow rate range in different mixer diameters, named region B in this paper, in which the air mass flow rate can be boosted with the rocket mass flow rate and the specific impulse ratio is higher than the rocket specific impulse. Significantly, the RBCC design state should be maintained in region B to meet the operational requirements of engines under the precondition of an insufficient flight Mach number. Topics Kerosene, Combustion engine, Thermodynamic properties, Thermodynamic cycles, Spacecraft propulsion engines, Rocket engines, Combustion, Fluid mechanics, Fluid flows, Fluid mixing NOMENCLATURE Abbreviations AbbreviationsDABdiffusion and afterburningMRmixed regimeRBCCrocket-based combined-cycleSLSsea level staticSMCsimultaneous mixing and combustionSRsupersonic regimeSSRsaturated supersonic regime Variables VariablesAcross-sectional areaCPspecific heat capacity at constant pressureCrratio between specific heat capacity of two streamsDcross-sectional diameter or equivalent diameterErentrainment ratioFthrusthenthalpyHheighth*total enthalpyIspspecific impulseIsp′enhanced specific impulseMaMach numberṁmass flowPstatic pressureP*total pressurePcambient pressureqaerodynamic flow coefficientRgas constantSentropyTstatic temperatureT*total temperatureTrtotal temperature ratiouvelocityVvolumeβthe factor of the compound-flow indicatorγspecific heat ratioηppropulsive efficiencyλvelocity coefficientσrcompression ratioτcombustion total temperature ratioφfuel equivalence ratio Subscripts Subscripts0, 1, y, m, 2, 3, and 40, 1, y, m, 2, 3, and 4 sectionsa, b, and cdemarcation point of modesfflightjinjected fuel for afterburningmaxmaximum valueminminimum valuepprimary streamrrocketssecondary stream I. INTRODUCTION The rocket-based combined-cycle (RBCC) engine is always deemed as a desirable scheme, which makes affordable and reusable space vehicles possible,1,2 because it bridges both the advantages of high Isp (specific impulse) of the air-breathing engine and a high thrust-to-weight ratio of the rocket engine.3 The ejector mode is the paramount working mode of the RBCC engine. The embedded rocket draws airflow into the engine from the surroundings to supply the secondary combustor with a pressurized airflow. RBCC-powered vehicles require only a 20% takeoff thrust compared to conventional rockets,1 making them liable to supersede a rocket booster. For nearly 70 years, a considerable amount of research has been devoted to the study of the RBCC engine. The ejector mode is crucial to the RBCC propulsion system and is considered a key research focus.4 Many important experiments have proved the feasibility of the RBCC ejector mode. In 1994, a series of ground tests at a Mach number range of 0–2.2 were carried out at TNO Prins Maurits Laboratory to study the operating principles of the RBCC engine ejector mode.5 Under the sea level static (SLS) condition, the thrust augmentation of the ejector mode was about 10%–20%, and at a Mach number of 2.2, the Isp augmentation is 2.3 times that of the rocket-only operation. Some key technologies involved in the ejector mode are efficient ejection technology, efficient mixing and combustion technologies, etc.4,6–10 The RBCC engine commonly operates in an extremely wide flight envelope at a very fast climb speed, so flow choking in the RBCC inner flow passage is very complicated. In addition, the flow choking status is also closely related to the operation of the embedded rocket.11 Gist et al.,12 Foster et al.,13 and Deturris14, et al. analyzed the flow patterns and the corresponding shear layer interactions inside the RBCC engine by conducting experimental investigations on the unique Fabri choke. Kanda et al.,15 Tomioka et al.,16 Tani et al.,17 and Lu18 et al. also indicated that a reasonable match between the rocket jet and the ejected air becomes an important technical premise for the efficient operation of the RBCC engine in a wide flight range. The flow choking status not only determines the backpressure resistance of the inlet but also affects the entrainment ratio and even has a chain reaction to the thrust and specific impulse performance. However, flow choking and thrust performance are not closely related in the research process. Therefore, only little information about the relationship between flow choking and thrust performance can be acquired. The one-dimensional analysis method is a very effective tool, which is conducive to ascertaining the influence rules between parameters and furnishing the direction for the research. Thus, a lot of modeling studies have been conducted on RBCC engines, including theoretical and experimental model analysis.19–31 Yang et al.32 established a one-dimensional model for the maximum thrust of the RBCC engine, realizing that the engine thrust was boosted with the entrainment ratio, which increased to the maximum but then decreased under specific rocket operating conditions (such as Mach number, flow rate, total pressure, total temperature, etc.). It could be observed from the findings that the thrust under the ejector mode relies both on the entrained air mass flow rate and the engine thermal efficiency. However, only one flow choking state was taken into consideration in this model, bringing inconsistency with the wide working range of the RBCC engine. Pastrone33 defined a criterion for judging the flow choking state using the values of the ejector-to-primary-nozzle-throat area ratio and the primary-to-secondary total pressure ratio (Pp*/Ps*). However, only subsonic solutions of the mixed streams were taken into account in the model. The hypothesis is not reasonable because when the rocket mass flow rate (mp) is large, the mixed stream is prone to be in a supersonic state. Kanda and Kudo34 proposed a one-dimensional calculation method to analyze the flow choking state. The isentropic condition is only available at the beginning of the interaction.35 Therefore, the model has higher accuracy than other models to solve the mixing problem under pressure mismatch conditions. However, the effect of flow choking on thrust has not been analyzed by the model. As mentioned above, although there are many analytical models on the RBCC engine, the relationship between the flow choking state and the thrust performance is undetermined. The thrust of the RBCC engine with the ejector mode is detected in some models, while the discussion of flow choking is simplified simultaneously. Among all combustion cycles in the ejector mode,6 the DAB (diffusion and afterburning) cycle has the highest thermal efficiency. It has a 30% augment in thrust compared with the SMC (simultaneous mixing and combustion) cycle during the take-off stage.1 Therefore, the DAB cycle is the most likely way to achieve better performance for the RBCC ejector mode. In the DAB cycle, the two processes of mixing and secondary combustion are separated from each other, showing benefit to the thermodynamic analysis. Therefore, to effectively promote the thrust performance evaluation under different flow choking conditions, the thermodynamic process of the DAB cycle in the RBCC ejector mode is analyzed, and a quasi-one-dimensional model is constructed in this paper. The energy conversion, performance trend, and influence factors are analyzed by thermodynamic cycle diagrams. The variable specific heat ratio is determined, and a unified aerodynamic choking criterion is established to systematically evaluate the engine performance in the one-dimensional model. The main purpose of this paper is to illustrate the relationship between the flow choking state and the thrust performance and then manifest the main mechanism of the rocket engine to the RBCC engine thrust. The results can be used to provide theoretical guidance to the geometry design and flight trajectory design. II. THE DEFINITION OF CHOKING IN THE RBCC EJECTOR MODE A simple RBCC engine schematic is illustrated in Fig. 1. The engine mainly comprises the inlet, embedded rocket, mixer, diffuser, secondary combustor, and nozzle. According to previous studies, the rocket plume and the inflow air have been named the primary and secondary streams, respectively. Variable indices are associated with the following engine flow path region: 0 at the inlet entry; 1 and m at the beginning and end of the mixer, respectively; 2 and 3 at the entrance and the exit of the secondary combustor, respectively; and 4 at the nozzle exit section. Besides, the RBCC engine has a wide working range of the ejector mode, and the pressure matching between two streams cannot be achieved in section 1. Therefore, it is assumed that there is a y-section where the pressure is matched between two streams. FIG. 1.View largeDownload slideSchematic and nomenclature of the RBCC engine.FIG. 1.View largeDownload slideSchematic and nomenclature of the RBCC engine. Close modal With the variation in the flow parameters of flight conditions, especially with the difference in the total pressure ratio (Pp*/Ps0*), the flow patterns are different in the engine. There are several kinds of choking in the internal flow passage of the engine. It can be divided into geometric throat, thermal choking, and flow choking. Normally, there are two geometric throats, namely, the inlet throat and the nozzle throat. If the fuel equivalent ratio of the combustor is high enough, thermal choking may appear. Flow choking is a unique phenomenon in the RBCC ejector mode compared with other air-breathing engines. According to the location of flow choking, the operating mechanism of the RBCC engine can be classified into three categories, namely, the saturated supersonic regime (SSR), supersonic regime (SR), and mixed regime (MR).32 For the SSR, secondary stream choking before the entrance of the mixer can be named flow choking Ⅰ. The inlet is simplified as an equal area channel, and the flow loss in the inlet is ignored in this paper, so the location of this choking is situated at the exit of the inlet. For the SR, the aerodynamic throat is formed somewhere in the mixer, and the choking can be named flow choking II. For the MR, flow choking is formed at the mixer exit, which can be named flow choking III. III. IDEAL THERMODYNAMIC CYCLE ANALYSIS OF THE RBCC WITH THE ENGINE EJECTOR MODE A. Assumptions of the thermodynamic cycle analysis The energy transfer process of the RBCC engine ejector mode is analyzed by the thermodynamic cycle. It is essential to reveal the influence factors and understand the ejector mode process with an ejector and afterburning. Some assumptions are made for simplification as follows:The flow in the inlet, primary rocket nozzle, and exhaust nozzle of the RBCC engine is isentropic.The pressure between two streams is matched at section y, and the parameters are uniform at the exit of the mixer.The process of combustion is deemed as constant static pressure. B. The thermodynamic cycle process analysis The approach for estimating the performance of the RBCC engine is based on thermal engine closed cycle analysis. As shown in Fig. 2, an ideal cycle is plotted on a thermodynamic P–V graph, in which the power capability variations in two streams in the ejector mode are illustrated in the diagram so as to clarify and understand the process. A thermodynamic T–S graph is also shown in Fig. 3, in which the thermal transformation of the whole system is clearly delineated. In these two diagrams, the section is marked in Fig. 1, and the superscript is used to distinguish the secondary flow from the primary flow. FIG. 2.View largeDownload slideThermodynamic P–V graph of the RBCC engine ejector mode.FIG. 2.View largeDownload slideThermodynamic P–V graph of the RBCC engine ejector mode. Close modal FIG. 3.View largeDownload slideThermodynamic T–S graph of the RBCC engine ejector mode.FIG. 3.View largeDownload slideThermodynamic T–S graph of the RBCC engine ejector mode. Close modal A rocket engine without injection and afterburning is deemed as the benchmark. As shown in Figs. 2 and 3, the rocket gas undergoes an ideal combustion and expansion process from state 0 to state 1. At the exit of the nozzle, the heated gas expands ideally to the atmospheric pressure (Pc). Point 1 to 0 is used to close the thermodynamic cycle. It is an ideal isobaric process. Hence, the cycle process of the rocket alone is 0-1-0. In the RBCC engine ejector mode, under the effect of the rocket, the secondary stream state is varied from point 0′ to 1′ during the ejecting process. The rock plume serves as the driving force to eject the secondary stream and strengthen the secondary stream energy. The two streams are the kinds of pressure matched at section y. During the mixing process, the primary flow transfers momentum and heat to the secondary flow, and the two streams are blended uniformly at the exit of the mixer. During the secondary combustion process, the pressurized mixed stream is burned in the secondary combustor and undergoes the heating and expansion process under the constant pressure condition. In the end, under the assumption of isentropic expansion, the mixed stream is expanded to the ambient pressure and then returns to the initial state in the ideal isobaric process. The cycle process of the primary stream is 0-1-y-m-3-4-0, and that of the secondary stream one is 0′-1′-y′-m′-3′-4′-0′. C. The energy conversion analysis and the influence factors Before the combustion stage, the secondary flow mainly increases its thermal energy by absorbing heat from the primary stream, and at the subsequent stage of mixing, the pressure increases remarkably. At the same time, the energy of the primary stream is lost. During the combustion process, the mixed stream is heated and expanded. Hence, the thrust performance of the RBCC engine is based on the secondary combustion effect. The energy consumed in the ejection and mixing processes is complemented during the secondary combustion. The thrust augmentation depends on the entrainment ratio. A remarkable thrust gain can be obtained when the air mass flow rate is large enough. IV. MATHEMATICAL MODELING FOR THE RBCC EJECTOR MODE A. Assumptions used in the quasi-one-dimensional model The inner flow of the RBCC ejector mode is a one-dimensional, adiabatic, frictionless, and steady flow.The primary stream and the free-stream conditions are supplied. Besides, the mixer is simplified as an equal area of a straight duct.The inlet is simplified as an equal area channel, and the flow loss in the inlet is ignored. Besides, the inlet area (As) is not fixed, and As ≥ Asy.The pressures of two streams are matched at section y. Psy = Ppy between section 1 and section y, but the primary and secondary streams do not blend.The process of the primary stream from section 1 to section y and the process of the mixed stream in the diffuser are regarded as isentropic reversible processes.The process of the secondary combustion is deemed as constant static pressure. B. The definition of some important parameters There are various parameters that affect the performance characteristics of the engine. The entrainment ratio (Er) and the compression ratio (σr) are vital parameters to appraise the ejecting and mixing performance, which are defined in Eqs. (1) and (2). The ratio between specific heat capacity of two streams (Cr) and the total temperature ratio (Tr), listed in Eqs. (3) and (4), respectively, is a crucial inlet parameter,Er=ṁsṁp,(1)σr=Pm*Ps*,(2)Cr=CPsCPp,(3)Tr=Ts*Tp*.(4) In this paper, a gaseous oxygen/liquid kerosene propellant rocket engine is applied as a primary stream supply system. The rocket gas components can be obtained by the calculation of the chemical equilibrium compositions,36 and then CP (specific heat capacity at constant pressure) and γ (specific heat ratio) can be obtained from the thermochemical tables.37 CP and γ at each section are obtained via iteration. With Tp (primary stream temperature) representing the temperature of the rocket plume in the range from 300 to 3500 K, CPp and γp can be identified using the following polynomial formula:CPp=959+0.79Tp−2.89Tp2104+5.33Tp3108−3.93Tp41012,(5)γp=1.415−3.64Tp104+3.10Tp2107−1.65Tp31010+5.28Tp41014−9.23Tp51018+6.71Tp61022.(6) CPs and γs in the Ts (secondary stream temperature) range from 300 to 1000 K are listed in Eqs. (7) and (8), and CPs and γs in the Ts range from 1000 to 3000 K are obtained using Eqs. (9) and (8), respectively, where Rs is a gas constant of the secondary stream,CPs=3.65−1.34Ts103+3.29Ts2106−1.91Ts3109+0.28Ts41012×Rs,(7)γs=1.369+2.92Ts104−8.19Ts2107+7.25Ts31010−2.04Ts41013−8.18Tp51017+7.38Tp61020−1.93Tp71023+1.78Tp81027,(8)CPs=3.04−1.34Ts103−0.49Ts2106+0.086Ts3109+0.0057Ts41012×Rs.(9) There is an important aerodynamic function Z(λ) for depicting fluid impulse, which could be derived from the momentum conservation equation changing to the form of an impulse function, which can be written as follows:38 ṁv + PA = cγṁZ(λ),(10)cγ=γ+12γ2γRT*γ+1,(11)Zλ=λ+1λ,(12)λ2=γ+1Ma22+γ−1Ma2,(13)where cγ and Z(λ) are significant aerodynamic functions for accounting the fluid impulse. λ is the velocity coefficient. The property of compound-compressible flow is differentiated by the value of a compound-flow indicator β,36,39β=Apy/Amγp1Mapy2−1+Asy/Amγs1Masy2−1.(14) As β < 0, two streams as a whole behave supersonically, which are referred to as the conditions of “compound-supersonic flow.” β = 0 corresponds to the “compound-sonic flow,” and β > 0 represents the “compound-subsonic flow.” C. Ejecting and mixing processes In the evaluation of ejecting and mixing performance, the parameters of the rocket, the stagnation parameters of the secondary flow, and the diameter of the straight mixer are all provided. The purpose of performance calculation is to determine the mass flow rate of the secondary stream and the average parameters of the mixed stream. The mixed flow properties at the exit of the mixing duct can be obtained by the conservation of mass, momentum, energy, and equation of state, which are shown below. The conservation of mass, energy, and momentum equations are as follows:ṁp + ṁs = ṁm,(15)ṁpCPpTp* + ṁsCPsTs* = ṁmCPmTm*,(16)cγpyṁpyZλpy+cγsyṁsyZλsy=cγmṁmZλm.(17) The physical parameters of the completely mixed stream can be derived from the parameters of the primary and secondary streams. The specific heat at the constant pressure of the mixed gas can be calculated using the following equation:ṁpCPp + ṁsCPs = ṁmCPm.(18) Equations (1), (3), (5), and (18) can be combined to yield an equation for CPm, namely,CPm=CPp1+ErCr1+Er.(19) Equations (1), (3), (4), and (16) can be combined to yield an equation for Tm*, namely,Tm*=Tp*1+ErCrTr1+ErCr.(20) Similarly, Rm and γm can be calculated using Eqs. (21) and (22), respectively,Rm=Rp + ErRs1+Er,(21)γm=γs1+ErCrγsγp+ErCr.(22) D. Flow choking state After the primary stream and the free-stream conditions are supplied, there is a definite value of the mass flow rate of the secondary stream for the specified Am. As λsy = 0, it means ṁs=0. The limitation of As to the secondary stream mass flow rate is broken by the assumptions of As ≥ Asy. It is a benefit to integrate flow choking Ⅰ with flow choking II for model analysis. If λsy = 1, the choking state of the inner flow can be found in flow choking Ⅰ. As a consequence, the relationship between ṁs and Z(λm) is provided in Eq. (13). Because of the Z(λm) function characteristic, ṁs ought to satisfy the following equations:Zλm=cγpyṁpyZλpy+cγsyṁsyZλsycγmṁm=fṁs≥2.(23) When the maximum value appears at λsy < 1 and Z(λm) > 2, the choking state pertains to flow choking II. In addition, if the maximum value appears at λsy < 1 and Z(λm) = 2, the choking state pertains to flow choking III. Therefore, by varying λsy from 0 to 1, all kinds of choking conditions during the mixing process can be studied. By comparing with the value of λsy and Z(λm) at ṁs=ṁsmax, the flow choking state (Ⅰ, II, and III) can be obtained as follows:λsy=1,Zλm>2,(24)λsy<1,Zλm>2,(25)λsy<1,Zλm=2.(26) Furthermore, Z(λm) > 2 corresponds to two values of λm, the subsonic one and the supersonic one. The value is determined by β [Eq. (14)]. If β < 0, λm > 1, and if β > 0, λm < 1. E. Process in the diffuser Aoki et al.35 argued that the primary and secondary streams mixed well through the pseudo-shock. In the ejector mode, the mixed streams enter the combustor at a subsonic speed. As λm ≤ 1, under the isentropic reversible assumption, λ2 at the exit of the expansion section can be obtained by the conservation of mass equation,qλmAm=qλ2A2.(27) As λm > 1, the mixed stream is a supersonic stream at the exit of the mixer. The supersonic flow would form a positive shock wave at the inlet of the diffuser section. As a result, when Z(λm) > 2 and β < 0, the shock loss would be generated. λ2 at the exit of the expansion section can be obtained as follows:q1λmAm=qλ2A2.(28) F. Secondary combustion process The fuel equivalence ratio (φ) denotes the amount of heating addition. Application of the steady flow energy equation to the combustor givesṁs + ṁph2*+ṁjhj=ṁs1+Er/Er+ṁjh3*.(29) Considering ṁs1 + Er/Er≫ṁj, then,ṁjhj=ṁs1+ErCp2T2*T3*T2*−1/Er.(30) The combustion total temperature ratio (τ) is given asτ=T3*Tm*(31)so that τ can be expressed asτ=φErhpr16.191+ErCP2T2*+1,(32)where hj is the heat of kerosene combustion, which is taken to be 43.2 MJ/kg, and the fuel equivalence ratio (φ) is varied from 0 to 1. In this paper, we assumed that the oxidant of secondary air could completely react with the injected fuel (ṁj), so φ = 1. When the mixed streams flow after the combustion, Ma3 and P3* are given as follows:40 Ma3=Ma2τ1+γ3−12Ma22−γ3−12Ma22,(33)P3*=Pm*1+γ3−12Ma221−1τγ3/γ3−1.(34) G. Expansion process in the nozzle The expansion process is isentropic, and the exhaust gas expands to the ambient pressure (Pc). Thus the nozzle exit Mach number (Ma4) and velocity (v4) asMa4=P3*Pcγ4−1/γ4−1γ4−12,(35)v4=γ4R4T4*1+γ4−12Ma42⋅Ma4.(36) From the previous study,32 the thrust of the RBCC is given asF=(ṁp+ṁs+ṁj)v4−ṁsvs.(37) The specific impulse (Isp) of the RBCC engine is denoted asIsp=Fṁj+ṁp.(38) The propulsive efficiency ηp is represented asηp=ṁp+ṁs+ṁjv4−ṁsvs⋅vsṁp+ṁs+ṁj⋅v422−ms⋅vs22.(39) V. RESULTS AND DISCUSSION Applying the model mentioned above, the thrust performance of an RBCC engine can be calculated under inlet conditions. It is essential to clarify the flight conditions of the RBCC engine with the ejector mode before identifying its thrust performance. On the one hand, the ejector mode is involved in subsonic, transonic, and supersonic speed domains, implying a very large speed range. Different flight conditions are responded by the inlet conditions and back pressure conditions of diverse engines, showing a great impact on the inner flow state. On the other hand, the primary flow and the secondary flow are commixtures in the inner flow passage and form flow choking. Flow choking is an important phenomenon in the RBCC ejector mode. It can not only make the engine overcome the adverse effects of fluctuating back pressure in the combustion chamber but also limit the air mass flow rate. The flow choking state will also be varied with different flight states. Table I illustrates four representative flight conditions that have been chosen, corresponding to four flight conditions, which are low Mach low altitude, low Mach high altitude, high Mach low altitude, and high Mach high altitude. In the calculation, Pp* and Tp* are fixed at 3 MPa and 3500 K, respectively. The interaction among the flight conditions, the diameter of the mixer, and the rocket flow are discussed. The theoretical specific impulse of the rocket (Ispr) can be obtained under the corresponding flight conditions, and the values are also given in Table I. In these cases, ṁp ranges from 0 to 4 kg/s, and Dm ranges from 160 to 200 mm. By integrating all the state point data, we can get the engine performance data under all types of Dm with the change in ṁp. TABLE I.Flight conditions and a theoretical specific impulse of the rocket.
. Hf (km)
. Maf
. Pf* (Pa)
. Pf (Pa)
. Ispr (s)
. Case 1 2 0.6 101 420 79 501 240 Case 2 6 0.6 60 238 47 218 251 Case 3 12s 2.0 151 810 19 399 266 Case 4 18 2.0 59 200 7 565 276
. Hf (km)
. Maf
. Pf* (Pa)
. Pf (Pa)
. Ispr (s)
. Case 1 2 0.6 101 420 79 501 240 Case 2 6 0.6 60 238 47 218 251 Case 3 12s 2.0 151 810 19 399 266 Case 4 18 2.0 59 200 7 565 276 View Large A. Low Mach and low altitude flight conditions The Isp contour diagram of case 1 is shown in Fig. 4, in which the yellow line implies that Isp is equal to Ispr and the critical ṁp represents ṁpc corresponding to each Dm. As can be observed from the diagram, Isp decreases with the increase in ṁp at different types of Dm. At the same ṁp, the larger Dm is a benefit to the Isp performance. Figure 5 shows the ṁs contour. It can be found that there is a sharp change corner on the ṁs isoline, and the corner of ṁp is defined as ṁpb. As ṁp>ṁpb, ṁs decreases sharply in each Dm. This is attributed to a sudden change in the state of flow choking in the mixer. When ṁp is low, flow choking III is formed in the mixer. With the increase in ṁp, the choking type in the mixer changes to flow choking II. The boundary line between flow choking II and flow choking III can be obtained by connecting the sharp change corners of different isolines. With the increase in ṁp, ṁs decreases sharply at first and then increases slightly. When ṁp rises above ṁpb corresponding to the flow choking boundary, ṁs drops rapidly, and then the ratio decreases linearly. FIG. 4.View largeDownload slideIsp contour of case 1.FIG. 4.View largeDownload slideIsp contour of case 1. Close modal FIG. 5.View largeDownload slideṁs contour of case 1.FIG. 5.View largeDownload slideṁs contour of case 1. Close modal Figure 6 reveals the RBCC engine thrust contour diagram. It can be seen from the diagram that the thrust trend is also changed significantly at the flow choking boundary. This is a chain reaction caused by the change in flow choking in the mixer. To evaluate the contribution of fuel for afterburning to the thrust, enhanced specific impulse Isp′=dF/dṁj is supplied. Figure 7 shows a contour map of Isp′. Based on the Isp′=Ispr relationship, two solid purple lines are labeled in the figure. The ṁp between two solid-line regions reveals that Isp′>Ispr. With the increase in ṁp, Isp′ continues to increase and attains the maximum value at the flow choking boundary; then Isp′ sharply increases and maintains the downward trend. The greater the value of Isp′, the better the effect of afterburning fuel on the thrust gain of the RBCC engine, and Isp′ FIG. 6.View largeDownload slideF contour of case 1.FIG. 6.View largeDownload slideF contour of case 1. Close modal FIG. 7.View largeDownload slideIsp′ contour of case 1.FIG. 7.View largeDownload slideIsp′ contour of case 1. Close modal The RBCC engine’s ṁp–Dm relationship can be subdivided into four regions in light of the characteristics of the Isp′ contour. The region bounded by the boundary of Isp′ is clarified as region B, where the left boundary corresponding to each Dm is ṁpa and the right boundary is ṁpb. The region of ṁp from 0 kg/s to ṁpa is region A. The region of ṁpb to ṁpc is region C, and the remaining region is region D. Region B is delineated in Figs. 2 and 3. It can be seen that the change rate of ṁs and thrust is comparatively small in region B. The changing trend of Isp′ and Isp with ṁp is plotted in Fig. 8. It can be found that although Isp decreases monotonously from its maximum value when boosting ṁp, Isp′ has a maximum value in ṁpb. Isp′ is larger than Ispr in region B, while Isp′ is lower than Ispr in the other regions. It is also indicated that as long as ṁp is less than ṁpc, the RBCC engine can capture the thrust gain. Nevertheless, as ṁpc>ṁp>ṁpb, the RBCC engine is not efficient. Two smaller RBCC engines with ṁp FIG. 8.View largeDownload slideChanging trend of Isp′ and Isp with ṁp.FIG. 8.View largeDownload slideChanging trend of Isp′ and Isp with ṁp. Close modal The P2* contour is described in Fig. 9. It can be observed that with the increase in ṁp, the pressurization effect of the primary stream and the combustion efficiency are reinforced. In order to assess the contribution of the rocket plume to the pressurization effect at different values of ṁp, a contour of dP2*/dṁp is plotted in Fig. 10. An obvious change is found at the boundary of flow choking. With the increase in ṁp, dP2*/dṁp increases rapidly at the initial stage, implying that the effect of pressurization is remarkable in the initial stage. When ṁp increases to a certain value, dP2*/dṁp increases slowly, and the dP2*/dṁp inflection point is reached at the flow choking boundary. When ṁp increases through the flow choking boundary, dP2*/dṁp has a very sharp rise, then gradually stabilizes, and increases in a similar linear trend. This phenomenon can be easily explained by the combination of the variation trend of ṁs in Fig. 5. In addition, it can also be concluded that the pressurization effect is not a direct factor affecting the thrust performance of the engine. FIG. 9.View largeDownload slideP2* contour of case 1.FIG. 9.View largeDownload slideP2* contour of case 1. Close modal FIG. 10.View largeDownload slidedP2*/dṁp contour of case 1.FIG. 10.View largeDownload slidedP2*/dṁp contour of case 1. Close modal Figure 11 manifests the propulsion efficiency of the engine. It can be seen from the diagram that the changing trend of the propulsion efficiency is almost the same as that of Isp. The propulsion efficiency diminishes with the increase in ṁp and enlarges with the increase in Dm. Hence, the Isp performance can be adopted to measure the changing trend of the engine propulsion efficiency. FIG. 11.View largeDownload slidePropulsive efficiency contour of case 1.FIG. 11.View largeDownload slidePropulsive efficiency contour of case 1. Close modal B. Low Mach and high altitude flight conditions Figure 12 illustrates a contour map of the Isp′ of case 2, with disparate regional divisions and flow choking ranges. The variation in Isp′ in diverse regions conforms to that of case 1, except that the range of flow choking III is significantly reduced. Figures 13–15 reveal the contour of the engine thrust, ṁs, and Isp, respectively. It can be found that with the increase in altitude, the thrust, ṁs, and Isp are reduced obviously, compared with the case 1 results. As the altitude of flight climbs to a higher level, the air total pressure diminishes. In case 1, the maximum value of Isp′ is 284 s, and that of Ispr is 240 s, so the gain is 18.3%; in case 2, the maximum value of Isp′ is 310 s, and Ispr is 251 s, so the gain is 23.5%. Thus, although the overall Isp decreases with increasing altitude, the benefits of the primary stream pressurization performance in region B increase and the advantages of the combined engine are more prominent. FIG. 12.View largeDownload slideIsp′ contour of case 2.FIG. 12.View largeDownload slideIsp′ contour of case 2. Close modal FIG. 13.View largeDownload slideF contour of case 2.FIG. 13.View largeDownload slideF contour of case 2. Close modal FIG. 14.View largeDownload slideṁs contour of case 2.FIG. 14.View largeDownload slideṁs contour of case 2. Close modal FIG. 15.View largeDownload slideIsp contour of case 2.FIG. 15.View largeDownload slideIsp contour of case 2. Close modal C. High Mach and low altitude flight conditions Figure 16 shows a contour map of the Isp′ of case 3, with different regions and flow choking ranges signified. It can be seen from the diagram that Isp′ is lower than 0 s when ṁp is low in region A. It implies that the energy of the rocket plume is completely consumed by the mixture. The rocket does not increase the thrust of the RBCC engine but causes a loss on it. Compared with the low Mach number condition, the ratio of region A to region B in flow choking III changes dramatically. With the growth of the Mach number, region B is compressed, and region A is dominant in the flow choking III region. Figure 17 delineates the P2* contour. The findings imply that the efficiency of the secondary combustion can be guaranteed by the higher total pressure at the inlet of the combustor. FIG. 16.View largeDownload slideIsp′ contour of case 3.FIG. 16.View largeDownload slideIsp′ contour of case 3. Close modal FIG. 17.View largeDownload slideP2*contour of case 3.FIG. 17.View largeDownload slideP2*contour of case 3. Close modal Figure 18 plots the Isp contour. It can be seen that the Isp of the RBCC engine is larger than that of the rocket alone within the range of calculation, showing that the rocket pressurization merely plays a supplementary role in such a flight condition while the thrust is generated in the RBCC engine, which takes full advantage of the entrained air. The thrust contour is plotted in Fig. 19, implying that the thrust first reduces but then increases with ṁp while such a phenomenon actually quite differs from the result under the low Mach flight condition because the rocket plume is required to meet the requirements of the mass flow and form sufficient thrust to compensate for the mixing loss during the pressurization process. The ṁs contour is described in Fig. 20. As ṁp increases, the growing trend of ṁs in flow choking III is not obvious, so the engine operating range can be set at region A or region B under high Mach and low altitude flight conditions. FIG. 18.View largeDownload slideIsp contour of case 3.FIG. 18.View largeDownload slideIsp contour of case 3. Close modal FIG. 19.View largeDownload slideF contour of case 3.FIG. 19.View largeDownload slideF contour of case 3. Close modal FIG. 20.View largeDownload slideṁs contour of case 3.FIG. 20.View largeDownload slideṁs contour of case 3. Close modal D. High Mach and high altitude flight conditions The contour of Isp′ under case 4 where disparate regions and aerodynamic choking ranges are manifested is plotted in Fig. 21. As the flight altitude increases, the total pressure is decreased. By making comparisons with the results in case 3, the negative thrust in region A is lowered, while the region of flow choking III is suppressed and the proportion of region A is reduced. In case 3, the maximum value of Isp′ is 300 s, and that of Isp is 266 s, with a gain of 12.8%; in case 4, the maximum value of Isp′ is 336 s, and that of Isp is 276 s, with a gain of 21.7%. Therefore, at a high Mach number, the ram effect of the engine will be reduced by increasing the altitude, while the contribution of the rocket pressurization to the engine thrust will be boosted. Figures 22 and 23 illustrate the thrust and ṁs contours, which are similar to the low Mach number situation. As the altitude increases, there is no doubt that the thrust and ṁs will be decreased dramatically, and the effective engine working state should be set to region B. FIG. 21.View largeDownload slideIsp′ contour of case 4.FIG. 21.View largeDownload slideIsp′ contour of case 4. Close modal FIG. 22.View largeDownload slideF contour of case 4.FIG. 22.View largeDownload slideF contour of case 4. Close modal FIG. 23.View largeDownload slideṁs contour of case 4.FIG. 23.View largeDownload slideṁs contour of case 4. Close modal Comparing Fig. 20 with Fig. 23 and Fig. 5 with Fig. 14, it can be found that when the air total pressure is small, the ṁs contour isoline has a parabola-like distribution in the flow choking III range. It means that ṁs diminishes first and then enlarges, and the growing trend is significant. On the contrary, if the total pressure of the air is large, the upward trend of ṁs in flow choking III is insignificant. In the mixer, as ṁp increases, the rocket plume will inevitably take up more space, resulting in compressing the air circulation area and cutting down the air flow. However, when the rocket flow rate is greater than a certain value in flow choking III, ṁs no longer decreases but instead increases. This indicates that the pressurization effect of the rocket plume on the air is remarkable under this condition. By raising the total pressure of the air, more air mass flow can be gained under the limited inner flow passage. In all cases, the region of increasing ṁs is shown in region B. This means that the higher performance in region B is conducive to the better pressurization of the rocket. E. Performance trend of the RBCC engine Based on the previous analysis, the ejector mode of the RBCC engine can be subdivided into two types of Ps*, one is lower Ps*, named working mode 1, and another is higher Ps*, called working mode 2. According to Fig. 24, The RBCC engine cannot organize the secondary combustion without the rocket pressurization at low Ps*. In region B, the air is pressurized enough, and the second combustion can be carried out. With the increase in ṁp, the secondary combustion efficiency is improved. If ṁp increases further, ṁs will decrease in region C. Although the pressurization effect is improved, Isp′ is less than Ispr. Even Isp is inferior to Ispr in region D. Under the circumstance, ṁs is crucial to the Isp performance of the RBCC engine. It is necessary to increase ṁs as much as possible for the high Isp performance. FIG. 24.View largeDownload slidePerformance trend of the RBCC engine in working mode 1.FIG. 24.View largeDownload slidePerformance trend of the RBCC engine in working mode 1. Close modal It can be observed from Fig. 25 that region A is further subdivided into region A1 and region A2 under the higher Ps* condition. In region A1, ṁp is small, and the mixing loss is superior to the rocket plume thrust contribution so that the RBCC engine cannot work at this stage. As ṁp increases further, the gain from the rocket pressurization effect is much higher than the loss from the mixing process. With higher Ps* air flow and the further pressurization by the rocket, the RBCC engine can be operated at region A2. At this point, the engine shows higher Isp performance. The contribution of engine thrust is made by higher Ps* air flow. To meet the requirements of the engine thrust, ṁp can be enlarged further. Both the rocket stream and the air stream provide more thrust in region B. The performance trend in region C and region D in working mode 2 is the same as that of working mode 1. The difference between these two working modes is mainly identified by Ps*. The performance of the RBCC engine in working mode 1 gets close to that of the rocket engine, while the RBCC engine in working mode 2 is more similar to the ramjet engine. FIG. 25.View largeDownload slidePerformance trend of the RBCC engine in working mode 2.FIG. 25.View largeDownload slidePerformance trend of the RBCC engine in working mode 2. Close modal VI. CONCLUSION The thermodynamic study was performed for analyzing the RBCC engine with the ejector mode. On the basis of the adiabatic, frictionless, and steady-flow hypotheses, a quasi-one-dimensional model was constructed to manifest the flow choking properties and the overall trend of thrust performance. The main conclusions could be concluded as stated below:With the increase in the rocket mass flow rate under all kinds of flight conditions, the flow choking state will be correspondingly varied, and there will be a sudden change in the engine performance during the flow choking transition period. The pressurization performance of the rocket cannot be adopted to represent the RBCC engine performance trend. The RBCC engine performance can be viewed as a composite of ramjet and rocket engines. The rocket engine contributes to the thrust by pressurizing the air, increasing the combustion efficiency of the combustion chamber. When the rocket mass flow rate is small, the effect of rocket pressurization is not enough to compensate for the loss caused by the mixing between the rocket plume and the air stream. When the mass flow rate of the rocket is large, the effect of pressurization is distinctive, but the rocket plume takes up too much space of the inner flow passage, resulting in a decreasing air mass flow rate of the air and the engine specific impulse. Consequently, the pressurization effect of the primary flow and the mass flow rate of the secondary flow ought to be incorporated together. Sufficient pressurization can guarantee the operation of the engine, and the key to achieving high specific impulse of RBCC engines is to ensure a higher air mass flow rate.The thrust of the RBCC engine stems from the pressurization effect of the rocket while the total pressure in the secondary stream is lower. On the contrary, the thrust is attributed to the airflow under the higher secondary stream total pressure condition. Besides, there is a critical value to rocket the mass flow rate under the higher Mach and low attitude flight conditions. If the mass flow rate of the rocket is inferior to the aforesaid value, the loss caused by the mixture between the primary and secondary streams cannot be supplemented by the pressurization effect of the rocket. Therefore, with the increase in the rocket mass flow rate, the engine thrust is reduced in this flight condition.Compared with flow choking II, the RBCC engine is more favorable in flow choking III. When the RBCC engine remains in flow choking III, the performance of the RBCC engine can be signified as RBCC engine ≥ ramjet + rocket. The RBCC engine integrates the advantages of the ramjet and the rocket engine under this condition. While the rocket mass flow rate increases further, the RBCC engine state is in flow choking II and region C. The performance of the RBCC engine can be denoted as [RBCC engine(smaller) + rocket] or [two RBCC engines(smaller)] ≥ RBCC engine ≥ ramjet + rocket. If the RBCC engine remains in region D, the performance of the RBCC engine is inferior to that of a rocket engine.In region B, the air mass flow rate can be increased with the rocket mass flow rate, and the specific impulse ratio is greater than the rocket specific impulse. When the flight Mach number is low, the RBCC engine should be in region B to guarantee the operation of the engine. As the Mach number increases, the rocket mass flow rate should be reduced gradually. At a higher Mach number, the RBCC engine ought to remain in region A to ensure the high specific impulse performance of the engine.Although the advantages of a combined engine are more prominent in the higher attitude, RBCC engines can generate more thrust at lower altitudes. The thrust augmentation of an RBCC engine is essentially due to a larger entrainment ratio. Therefore, the lower flight altitude should be chosen in the ejector mode so that the engine can gain enough acceleration, climb quickly, and enter the ramjet mode as soon as possible. ACKNOWLEDGMENTS This research work was supported by the National Natural Science Foundation of China (Grant No. 11902353). The authors would also like to express their sincere thanks for the constructive suggestions of reviewers in improving the quality of this paper. AUTHOR DECLARATIONS Conflict of Interest The authors have no conflicts to disclose. Author Contributions Yizhi Yao: Conceptualization (lead); Data curation (lead); Formal analysis (lead); Funding acquisition (lead); Visualization (lead); Writing – original draft (lead); Writing – review & editing (equal). Rui Gu: Data curation (supporting); Software (equal); Writing – review & editing (supporting). Mingbo Sun: Project administration (equal); Resources (equal); Writing – review & editing (equal). Peibo Li: Funding acquisition (equal); Project administration (equal); Resources (equal); Validation (equal). Yuhui Huang: Funding acquisition (supporting); Resources (supporting); Writing – review & editing (supporting). Bin An: Data curation (supporting); Formal analysis (supporting); Funding acquisition (supporting); Writing – review & editing (supporting). Jiaoru Wang: Formal analysis (supporting); Writing – original draft (supporting); Writing – review & editing (supporting). Menglei Li: Visualization (supporting); Writing – original draft (supporting); Writing – review & editing (supporting). Taiyu Wang: Writing – original draft (supporting); Writing – review & editing (supporting). Jikai Chen: Writing – review & editing (supporting). DATA AVAILABILITY The data that support the findings of this study are available from the corresponding authors upon reasonable request. REFERENCES 1.A. Siebenhaar, M. J. Bulman, and D. K. 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百年刊庆 清华大学学报（自然科学版） 2017, Vol. 57 Issue (5): 516-520 DOI: 10.16511/j.cnki.qhdxxb.2017.22.030
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组合循环推进系统燃料消耗模型及优化分析 计自飞, 王兵, 张会强 清华大学 航天航空学院, 北京 100084 Optimal analysis of the fuel consumption of combined cycle propulsion systems JI Zifei, WANG Bing, ZHANG Huiqiang School of Aerospace Engineering, Tsinghua University, Beijing 100084, China 摘要 图/表 参考文献 相关文章 Metrics 全文: PDF(1332 KB)
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摘要 根据飞行任务要求，准确计算出飞行器所需的燃料消耗是推进系统设计的前提。该文针对火箭基组合循环动力（RBCC）推进方式，并以“地面起飞—巡航—滑翔着陆”的高超音速飞行器为研究对象，采用理论分析的方法建立了燃料消耗的计算模型，并提出一种以最小燃料消耗为目标的多参数优化方法。周期跳跃式巡航飞行器燃料消耗的研究结果表明：随着巡航初速度、爬升段航迹倾角、巡航轨迹角的增加，燃料消耗量增加；随着飞行动压的增加，燃料消耗量先减小后阶跃式增加。优化分析结果表明：对于起飞质量100 t、2 h全球到达的RBCC组合动力高超音速飞行器，在升阻比为4时巡航跳跃周期数为46，最小燃料消耗量约为32 t。研究结果表明该燃料消耗分析方法合理、可行，为高超音速飞行器及组合循环动力推进系统的工程设计提供了依据。 服务 把本文推荐给朋友 加入引用管理器 E-mail Alert RSS 作者相关文章 计自飞 王兵 张会强 关键词 ：
组合循环推进,
高超音速飞行,
燃料消耗,
优化方法
Abstract：The propulsion system fuel consumption must be accurately predicted for aircraft missions. A theoretical analysis of a hypersonic aircraft with a “boost-cruise-glide” flight mission profile powered by a rocket based combined cycle (RBCC) engine is used to predict the fuel consumption of the aircraft and to optimize the fuel consumption. The fuel consumption analysis of periodic hypersonic cruise trajectories shows that the fuel consumption decreases with increasing initial cruise velocity, larger flight-path angles and larger flight-path angles for cruising. As the flight dynamic pressure increases, the fuel consumption first decreases but then increases with a step change. The optimization results show that the hypersonic cruise stage should have 46 skip-periods with a minimum fuel consumption of about 32 tons for a two-hour global-reaching hypersonic aircraft with an initial weight of 100 tons and a lift-drag ratio of 4. The optimal results show that the fuel consumption prediction model is reasonable. The present study can guide the design of combined cycle propulsion systems. Key words：
combined cycle propulsion
hypersonic flight
fuel consumption
optimization method 收稿日期: 2015-12-11
出版日期: 2017-05-15
ZTFLH: V43
通讯作者:
王兵,副教授,E-mail:wbing@tsinghua.edu.cn
E-mail: wbing@tsinghua.edu.cn
引用本文:
计自飞, 王兵, 张会强. 组合循环推进系统燃料消耗模型及优化分析[J]. 清华大学学报（自然科学版）, 2017, 57(5): 516-520.
JI Zifei, WANG Bing, ZHANG Huiqiang. Optimal analysis of the fuel consumption of combined cycle propulsion systems. Journal of Tsinghua University(Science and Technology), 2017, 57(5): 516-520.
链接本文:
http://jst.tsinghuajournals.com/CN/10.16511/j.cnki.qhdxxb.2017.22.030
或
http://jst.tsinghuajournals.com/CN/Y2017/V57/I5/516
图1 火箭置于中心锥体的一种RBCC结构示意图
图2 RBCC发动机比推力以及比冲随飞行Ma 数的变化
图3 优化方法流程图
表1 优化设计参数的限制条件
图4 燃料消耗、巡航周期数和总时间随巡航初速度的变化
图5 燃料消耗、巡航周期数和总时间随巡航轨迹角的变化
表2 以最省燃料为目标优化飞行参数
图6 RBCC动力飞行器飞行过程优化结果示意图
图7 3种推进方式的最小燃料消耗量对比
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组合循环推进系统燃料消耗模型及优化分析