基于内乘波概念的TBCC进气道过渡模态研究

TBCC(涡轮基组合循环)在过渡模态的工作,是这种新型高速动力系统的关键问题。本文针对基于内乘波概念的内并联式TBCC进气道进行了其过渡模态流场的研究。首先,应用本团队自主提出的内乘波式进气道设计概念,提出一种新型TBCC进气道设计方案。然后,验证了其在设计马赫数4.0下实现全流量捕获,在模态转换时,分流板的分流是有效的。针对其在过渡模态马赫数2.5的性能,通过基于数值模拟的设计参数分析,发现在内参考收缩比为3.11时进气道流通面积合适可实现自起动、且综合性能较优。研究还发现,分流板偏转角度范围越小,分流段流动越好、通流出口总压恢复系数也越高。     

一种新型的超燃冲压发动机闭式冷却循环

热防护技术是发展高超音速的关键技术之一.为了降低冷却用燃料需求,本文提出了基于闭式Brayton循环的超燃冲压发动机冷却循环,旨在提出一个缓解超燃冲压发动机热防护压力的新方法、新思路,同时可为高超音速飞行器提供足够的电力供应.文中介绍了冷却循环的组成和工作原理,定义了评价冷却循环性能的参数,在考虑系统内、外不可逆性的基础上,完成了系统的性能分析,分析表明该冷却循环可有效地节约冷却用燃料消耗,大幅提升超燃冲压发动机热防护系统性能.     

涡轮过渡段气动性能数值优化

采用CFD数值模拟方法对中档功率燃气轮机燃气涡轮与动力涡轮之间的涡轮过渡段进行气动性能优化设计.在原始涡轮过渡段流道及承力支板基础上,新设计进口导流叶片,优化过渡段流道,并采用Numeca软件Design3D 优化平台计算机自动寻优,数值模拟相对比较说明,新设计的涡轮过渡段气动性能满足指标要求.涡轮过渡段承接燃气涡轮、...     

考虑气蚀效应的气液两相冲压发动机数值模拟

针对一种新型水下气液两相冲压发动机,综合考虑了湍流效应、气液两相之间的拖曳作用及传热与传质,利用计算流体力学方法研究了气液两相冲压发动机内流场的流动特性随发动机工作条件的变化规律,重点研究了气蚀效应对发动机工作性能的影响.主要结论为:当航行速度大于32 m/s,气液两相冲压发动机入口附近会产生气蚀并造成严重的总压损失,导致扩张段下游产生流动分离,此时发动机无法产生正推力;通过增大气体质量流率,气液两相冲压发动机内流场的压力将会随之升高,气蚀效应被抑制;提高注入发动机气体的温度,发动机的推力及比冲均增大,但是发动机效率急剧降低.      

E3对转涡轮设计特点分析(Ⅱ)--过渡段不同方案的流场分析比较

本文以E3对转涡轮方案为背景,针对高压涡轮动叶设计难以满足低压涡轮预旋需求的难点,提出了取消低压涡轮导叶,改由过渡段支板弥补高压涡轮所提供预旋不足的过渡段改进设想.对过渡段支板不同改进方案和E3原方案进行了流场计算、气动性能比较及结构可行性分析.     

超燃冲压发动机推进性能理论分析

超燃冲压发动机的正推力问题和超声速燃烧的稳定性问题是制约超燃冲压发动机发展的两个关键气动物理问题.虽然经过50多年的研究,但是目前国内外对这两个关键问题的机理还没有研究清楚.文章首次将CJ爆轰理论应用于超燃冲压发动机推进性能分析,给出了这两个关键气动问题的理论分析结果.分析结果表明,燃烧室入口空气静温对发动机的推进性能产生重要影响.当爆轰波的爆速大于隔离段内空气来流的速度时,会向隔离段上游传播,导致发动机不起动.飞行Mach数Ma=6~8是超燃发动机的临界不稳定范围,飞行Mach数Ma>9,超声速燃烧将变得稳定.      

某型燃机涡轮过渡段流场优化设计

通过对某型燃机涡轮过渡段流场的优化设计提高了其气动性能.同时针对于这类带有导流支板的涡轮过渡段的优化设计总结出了如下规律:过渡段导流支板的转折角宜小不宜大;其安装角度应尽量接近轴向;过渡段子午流道形状及导流支板的外型要相互配合以满足亚音速扩压流道的要求,而这种逐渐扩压流道的气动性能最佳.     

旁通补燃动力涡轮复合式发动机的理论研究

本文在研究法国Hyperbar柴油机的基础上,提出了旁通补燃动力涡轮复合发动机方案。并对它的实际工作过程、性能特点、提高功率的潜在能力和配套方式进行了电算预测。从理论上证明了这种复合发动机方案不仅具有Hyperbar系统的一切优点,并且比Hyperbar具有更大的优越性。在高功率的移动式发动机方面,提出了一个新的概念和方向。计算发现了这种复合式发动机特有的最佳增压比和最佳涡轮前温度的概念,它对这种发动机的性能有重大的影响。在耗油率曲线上还发现存在着一种独特的凹弧。     

加力燃烧室两相混气燃烧研究

一、前言 在涡轮风扇发动机加力室外函和冲压发动机燃烧室中,由于气流温度低,燃油蒸发性能差,混气在稳定器截面仍是两相的。气流绕稳定器运动过程中,小油珠随气流运动。大油珠由于惯性大,撞落在稳定器表面,形成油膜。油膜沿壁面向稳定器后缘运动过程中吸热、升温和蒸发。蒸发的燃料随气流向下游运动和扩散。剩余未蒸发燃油在气流剪切力     

烧蚀支板对超燃冲压发动机燃烧室性能的影响

为了提高超燃冲压发动机燃烧室的性能,本文提出了燃料喷注支板与烧蚀支板组合的燃烧室新方案,并研究了新方案对超燃冲压发动机燃烧室性能的影响。相比于单燃料喷注支板方式而言,加入烧蚀支板后,虽然燃烧室内的总压恢复系数有所下降,但燃烧室内燃料与空气的混合效率、燃烧效率均有显著提高,燃烧效率的提高弥补了燃烧室内总压损失所带来的机械能损失,使得燃料喷注支板和烧蚀支板组合方式下的燃烧室比冲高于单燃料喷注支板时的比冲。     

用于高超音速推进的反循环发动机热力性能分析

符号表fIkg空气对应的燃料流量ISP燃料比冲Ti后燃烧室出口温度马比推力，燃料空气当量比WT涡轮输出功MT飞行马赫数＄前登燃烧室出口温度。T涡轮膨胀比H飞行高度T4＊热交换器出口温度7TC压气机压比1引言八十年代以来，美、俄、德、法、英、日等主要空间大国均提出了各自的高超音速计划，如NASP、Sanger、HOTOL、STAR等。各国的方案在Ma＞6．5均采用火箭发动机或超燃冲压发动机，对于Ma＜6．5采用何种方案则分歧较大，有涡轮一冲压组合发动机方案、ATR方案、LACE方案等。本文对反循环发动机（InverseCgcleEngine）方案作了进…     

预注冷质的空气涡轮火箭性能分析研究

提出了压气机进口预注冷质的空气涡轮火箭(MIPCC ATR)循环,拓展了ATR在高飞行马赫数下的工作能力。给出了基于工质组分的变比热ATR循环热力学建模方法和工质物性改变时对部件特性的修正方法,基于所推导的准无量纲关系式讨论了循环参数之间的相互作用及其对性能的影响。研究结果表明预注冷质不仅能有效拓宽ATR的工作包线,而且可以提高单位推力,所需付出的代价是比冲会有一定程度的降低。     

超声速气流中的爆震推进理论与研究进展

爆震燃烧近似为等容燃烧,理论上其热循环效率高于基于等压燃烧的爆燃燃烧,在超声速推进系统中具有潜在的应用价值.通过总结超声速气流中的爆震推进理论与研究进展,分析其需要解决的关键科学与技术问题,指导未来高超声速发动机的基础研究.文章重点总结了适用于高超声速飞行的斜爆震发动机、超声速脉冲爆震冲压发动机的基础研究进展.其中对斜爆震发动机的应用模式、相关实验研究思路及方法、数值仿真现状进行了总结分析.对超声速脉冲爆震冲压发动机的基础理论研究现状和目前研究的难点进行了梳理.基于爆震燃烧的超燃冲压发动机具有推进系统自增压、燃烧效率高、推力性能好、推进效率高、燃烧室长度短、结构重量轻等优势,文章总结了该发动机当前的发展进程和最新的研究进展,并对其未来的发展方向以及存在的技术问题进行了分析.      

Thrust characteristics of an airbreathing pulse detonation engine in supersonic flight at various altitudes

The main thrust characteristics, such as thrust force, specific impulse, specific fuel consumption, and specific thrust, of a pulse detonation engine (PDE) with an air intake and nozzle in conditions of flight at a Mach number of 3 and various altitudes (from 8 to 28 km above sea level) are for the first time calculated with consideration given to the physicochemical characteristics of the oxidation and combustion of hydro-carbon fuel (propane), finite time of turbulent flame acceleration, and deflagration-to-detonation transition (DDT). In addition, a parametric analysis of the influence of the operation mode and design parameters of the PDE on its thrust characteristics in flight at a Mach number of 3 and an altitude of 16 km is performed, and the characteristics of engines with direct initiation of detonation and fast deflagration are compared. It is shown that a PDE of this design greatly exceeds an ideal ramjet engine in specific thrust, whereas regarding the specific impulse and specific fuel consumption, it is not inferior to the ideal ramjet.     

Flame dynamics under various backpressures in a model scramjet with and without a cavity flameholder

Flame dynamics under various backpressure conditions were experimentally investigated using direct flame visualization, high-speed CH* chemiluminescence imaging, and wall pressure measurements. The stagnation pressure and temperature used in the present study were 100 kPa and 2500 K, respectively, with a freestream Mach number of 4.5. Rectangular scramjet models with and without a cavity were used to explore the effects of the cavity on flame dynamics when operating in scramjet mode, ramjet mode, and unstart. The flow rate of the ethylene jet was varied to impose backpressures corresponding to each operation mode. For both models, reverse flame propagation was observed for ramjet mode and unstart. For ramjet mode, flame fluctuation occurred within the isolator due to the coupling of fluid dynamics and combustion. The presence of a cavity enhanced combustion and reduced flame fluctuation in both scram and ramjet mode. The cavity promoted unstart because of the greater heat release from combustion. Further research using spatially resolved diagnostic techniques is needed to identify the flame locations for ramjet mode and unstart.     

双余度发动机状态监测技术的研究与应用

在飞机、舰艇、装甲、汽车电子控制系统中采用双余度发动机状态监测技术,能够有效提高飞机、舰艇、装甲、汽车运行的安全性和可靠性,以某型涡轴发动机为研究对象,对双余度发动机状态监测技术进行了深入研究,研制了双余度发动机状态监测系统,介绍了系统的总体设计、软硬件设计、余度设计,虚拟仪表设计,并在双余度发动机状态监测系统的设计中引入了嵌入式PC/104模块;为了使状态监测系统具有更好的扩展性和适应性,将系统设计成可以在两种方式下工作:机载运行方式和地面试车方式,通过这两种方式对发动机状态进行自动监测,为发动机的状态趋势分析、故障诊断和视情维修提供科学的依据;采用双余度发动机状态监测技术研制的某型涡轴发动机状态实时监测系统经过了大量的地面试验和某型直升机上试飞试验,功能、性能满足要求,目前该系统已在某型直升机上得到应用。     

Exergy method for estimating the ramjet specific impulse

The exergy method is developed for computing the ramjet thrust-economic characteristics with regard for real thermodynamic properties of combustion products when using as fuel the hydrogen and hydrocarbon fuel for the freestream Mach numbers M = 4 ÷ 14. The estimates for the specific impulse of the given engine using the presented technique are shown to agree with the estimates computed by other authors. The computational method is intended for obtaining the ramjet characteristics and conduction of the parameter analysis at the research initial stage as well as for its use at the conceptual developments of hypersonic flying vehicles.     

Thrust characteristics of an airbreathing pulse detonation engine in flight at mach numbers of 0.4 to 5.0

Multidimensional calculations are performed to demonstrate that, by its characteristics, the pulse detonation engine (PDE) is a unique type of ramjet propulsion system, which can be used in both subsonic and supersonic aircraft. By a number of examples, it is shown that, in various thrust characteristics, such as the specific impulse, specific fuel consumption, and specific thrust, the PDE substantially exceeds ramjet engines.     

组合式发动机进气道的冷却研究

针对空天飞机组合发动机进气道的冷却要求，采用冲击－对流复合冷却技术对进气道进行强化冷却试验，按照得出的冷却结构各部分传热经验公式，用温度200K，压力1MPa低温氢作冷却剂，算出进气道的冷却热流密度，证明所用结构可以满足进气道冷却要求。     

Estimation of energetic efficiency of heat supply in front of the aircraft at supersonic accelerated flight. Part 1. Mathematical models

Fuel economy at boost trajectory of the aerospace plane was estimated during energy supply to the free stream. Initial and final flight velocities were specified. The model of a gliding flight above cold air in an infinite isobaric thermal wake was used. The fuel consumption rates were compared at optimal trajectory. The calculations were carried out using a combined power plant consisting of ramjet and liquid-propellant engine. An exergy model was built in the first part of the paper to estimate the ramjet thrust and specific impulse. A quadratic dependence on aerodynamic lift was used to estimate the aerodynamic drag of aircraft. The energy for flow heating was obtained at the expense of an equivalent reduction of the exergy of combustion products. The dependencies were obtained for increasing the range coefficient of cruise flight for different Mach numbers. The second part of the paper presents a mathematical model for the boost interval of the aircraft flight trajectory and the computational results for the reduction of fuel consumption at the boost trajectory for a given value of the energy supplied in front of the aircraft.