管路式环形散热器性能数值分析研究

基于高隐身飞行器对散热器的设计要求,提出一种安装在进气道末段的新型环形散热器,对其进行数值仿真模拟研究,得到了改型散热器改善的性能,对其进行了质量预估,并将计算结果与传统波纹板结构形式散热器和传统翅片结构形式散热器进行比对。计算结果表明,新型环形散热器较传统波纹板结构形式散热器和传统翅片结构形式散热器相比,新型环形散热器性能明显优于传统型散热器,其换热密度平均提升379%、气体侧换热压降比提升1 470%,即相同换热量时,新型换热器具有更轻的质量、更小的流阻。同时,新型环形散热器采用3D打印技术制造,一体化成型,耐压性更好,且具有快速试制、快速响应、产品一致性好等优点。     

航空发动机管壳式燃-滑油散热器换热特性计算

采用无量纲关系曲线(Re~St·Pr2/3)计算管壳式燃-滑油散热器管程、壳程表面传热系数,基于效率-传热单元数法建立了管壳式燃-滑油散热器换热特性计算模型;以航空发动机某管壳式燃-滑油散热器少量试验数据为基础,拟合得到了该型散热器管程和壳程Re~St·Pr2/3关系曲线,并将该关系曲线应用于另一结构类似的管壳式燃-滑油散热器换热特性计算.算例计算表明:计算值与试验数据吻合较好,管程、壳程出口温度相对误差均不超过6.5%,总换热量相对误差不超过7.4%.该方法可应用于航空发动机管壳式燃-滑油散热器设计及验证阶段的换热特性计算.      

管壳式燃滑油散热器换热特性

管壳式燃滑油散热器换热特性直接影响航空发动机滑油系统的散热冷却能力,对滑油系统的热平衡建立至关重要。为了更准确地进行滑油系统热分析计算,优化滑油系统设计,必须掌握更为精确的燃滑油散热器换热特性的计算方法。在对散热器壳程的复杂流动进行分析的基础上,分别采用Kern法、Bell-Delaware法和分段模拟法计算换热特性,并通过试验结果验证计算准确度,经对比表明,将壳程换热按流动特性进行分段模拟计算的方法具有更高的准确度,满足滑油系统设计需要。     

管壳式燃滑油散热器换热特性计算方法及试验验证

管壳式燃滑油散热器换热特性直接影响航空发动机滑油系统的散热冷却能力,对滑油系统的热平衡建立至关重要.为了更准确地进行滑油系统热分析计算,优化滑油系统设计,必须掌握更为精确的燃滑油散热器换热特性的计算方法.在对散热器壳程的复杂流动进行分析的基础上,分别采用Kern法、Bell-Delaware法和分段模拟法计算换热特性,并通过试验结果验证计算准确度,经对比表明,将壳程换热按流动特性进行分段模拟计算的方法具有更高的准确度,满足滑油系统设计需要.     

基于飞发一体化的滑油系统热性能仿真

航空发动机滑油系统与飞机、发动机的关联参数有限。为准确表达变工况滑油系统的热性能,通过研究发动机轴承腔热性能与转子转速及主流路温度参数的拟合关系,将主机温度、燃滑油参数作为输入,对发动机滑油系统在飞行剖面上典型飞行状态点的热性能参数进行了迭代计算;针对管壳式燃滑油散热器结构及运行特性,计算了散热器换热性能。建立轴承腔和散热器的数学模型;基于系统流动仿真平台,利用内部的二次开发环境编写出C#语言代码,开发出了适用于发动机的轴承生热模型和散热器模型,实现发动机滑油系统与发动机燃油系统及飞机热管理系统的联合计算;在航空发动机、飞机变工况输入条件下,进行滑油系统、发动机整机及飞发一体化的变工况热性能迭代计算,并与试验数据进行对比。结果表明：该计算方法误差小于5%,可较准确地反映变工况条件下的热管理相关参数,为飞发一体化热管理联合仿真分析提供可靠的数据来源。     

紧凑翅片管式空气-燃油换热器试验

为实现航空发动机内部热环境与热沉的有效交互，探究换热元件的流动换热特性。以航空发动机燃油系统回油冷却换热器为例，开展了小管径矩形翅片管式空气-燃油换热器流动换热性能试验研究。试验采用高温燃油与常温空气两股工质在换热器中进行能量交换，探究换热器在不同工况下的流动与换热性能，获取矩形翅片管式换热单元管外流动换热经验关联式。结果表明：矩形翅片管式换热单元的表面传热系数约为相同结构参数光滑管束换热单元的44%，且试验结构换热单元阻力系数高于光滑管束单元，在进行翅片管束换热器设计时应综合考虑翅片对流动换热性能的影响。试验获取的翅片管式换热单元管外努塞尔数经验关联式与阻力系数经验关联式拟合偏差均不超过5%，较为准确地反应了换热单元外侧的流动换热特性。     

环形散热器的动态性能

 利用热阻热容概念近似描述环形散热器微元段上的动态特性,加上热平衡条件,从而推导出整个环形散热器动态数学模型。只要在微元段上选取适当的节点数,该模型就可以满足任意精度之要求。借助于计算机并使用有限差分法,对环形散热器的动态性能进行了数值计算,理论计算结果与空中飞行试验数据吻合较好。还给出了环形散热器芯体在某一状态下的温度分布,作为动态响应计算的特例,最后计算了此状态下环形散热器稳态性能,其结果与对应的试飞数据比较令人非常满意。     

环境温度对散热器性能试验的影响

本文主要讨论航空空气——空气散热器的环境温度对其性能的影响。文中假设了环境对性能影响的各种极值条件,用定量推算法得出“环境温度对航空空气散热器性能影响很小,可以忽略不计”的结论,并用试验资料作为该结论的间接旁证。     

凹坑强化传热的研究进展回顾

基于涡流发生的凹坑强化换热是一种新型高效的强化换热和冷却技术,其主要特点是传热强度大,流动阻力小,综合传热性能高.通过近十年来凹坑强化换热研究工作的回顾,介绍凹坑壁面的换热阻力特性和涡流结构,并讨论凹坑形状、深度和雷诺数等参数的影响.通过介绍凹坑在旋转通道、圆管和冲击耙面等不同情况下的换热阻力特性,以及与肋等的综合传热性能对比,旨在说明凹坑是一种很具潜力的新型强化换热结构,具有十分广泛的应用前景.      

RWT800散热器风洞研制

散热器广泛应用于各种动力机械和仪器仪表，它的性能优劣直接影响产品性能。对于动力机械来说，特别关注其散热效率、噪声和成本。散热效率与能耗相关，噪声直接与人的健康相关。提高散热器性能必需有相应的研究、检验设备，散热器风洞就是重要设备之一。2002年，西北工业大学和恒安散热器有限公司联合研制了一座RWT800散热器风洞，该风洞指标先进，性能优良，是我国机械工业系统至今先进的散热器风洞。     

履带车辆前置动力舱冷却空气流场分析

以某履带车辆前置动力舱为研究对象,建立了舱内冷却空气流场的物理模型与数学模型,对装备有多个散热器且入口边界条件不同的动力舱内、外部冷却空气流场进行了仿真研究,并与实车测试结果进行了对比,分析了不同车速对各散热器冷却空气质量流量的变化规律的影响.结果表明:在相同风扇转速下,随着车速由0km/h增加到70km/h,两个散热器及两个风扇的质量流量均有10%左右的变化,但是两个散热器的总质量流量基本不变.车速在0~70km/h变化时,由仿真计算和试验测试所得的各散热器和风扇的冷却空气质量流量变化率之差都小于5%.     

ALE-9开阵天线辐射柱故障分析

针对单脉冲二次雷达所使用的ALE-9开阵天线,研究了辐射柱故障对雷达水平波束的影响。通过分析发现雷达天线各辐射柱的状态对天线水平波束的影响是不同的。通过大量的计算找出影响水平波束形状的关键柱体,从而在实际中加以重点监测和维护。     

美国用于空间站辐射器中的热控涂层

综述了六种美国可用于空间站辐射器中的热控涂层。从涂层的性能、价格、质量以及成熟性研究，对这六种热控涂层进行了评价。结果表明，Z－93型热控涂层最适合用于空间站的辐射器中，镀银F－46薄膜型热控涂层次之。     

The IRIDIUMR main mission antenna concept

The design of a novel phased array panel that provides the L-band satellite to ground links for the IRIDIUM global communications system is presented. Key components and aspects of this phased array antenna are discussed, including the beamforming architecture, radiated intermodulation products, the patch radiators, and the T/R module. The strategy for minimizing DC power consumption over a large range of multicarrier rf output power is described. Finally, test results showing compliant array operation are summarized     

Electron,Proton, and Alpha Monitor on the Advanced Composition Explorer spacecraft

The Electron, Proton, and Alpha Monitor (EPAM) is designed to make measurements of ions and electrons over a broad range of energy and intensity. Through five separate solid-state detector telescopes oriented so as to provide nearly full coverage of the unit-sphere, EPAM can uniquely distinguish ions (Ei≳50 keV) and electrons (Ee≳40 keV) providing the context for the measurements of the high sensitivity instruments on ACE. Using a ΔE×E telescope, the instrument can determine ion elemental abundances (E≳0.5 MeV nucl−1). The large angular coverage and high time resolution will serve to alert the other instruments on ACE of interesting anisotropic events. The experiment is controlled by a microprocessor-based data system, and the entire instrument has been reconfigured from the HI-SCALE instrument on the Ulysses spacecraft. Inflight calibration is achieved using a variety of radioactive sources mounted on the reclosable telescope covers. Besides the coarse (8 channel) ion and (4 channel) electron energy spectra, the instrument is also capable of providing energy spectra with 32 logarithmically spaced channels using a pulse-height-analyzer. The instrument, along with its mounting bracket and radiators weighs 11.8 kg and uses about 4.0 W of power. To demonstrate some of the capabilities of the instrument, some initial performance data are included from a solar energetic particle event in November 1997. This revised version was published online in June 2006 with corrections to the Cover Date.     

Hypersonic Re-Entry Optical Transfer Equations And Photographic Techniques

Radiation transfer equations applicable to various types of imaging instruments used against distant sources are presented. Emphasis is placed on measurements against point and line radiators made with instruments yielding image spectra. Framing and streaking cameras are discussed in terms of the overall transfer functions of the instrument and sensor. Calibration techniques used for absolute intensity measurements are shown to yield data within a factor of two of the actual value in most cases. The instruments described are of specific use in optical radiometry against bodies penetrating the earth's atmosphere in the hypervelocity regime, and are equally applicable to measurements in a ballistic range. For the problem of tracking these fast-moving objects, imaging instruments are preferred to point detector devices.     

City Noise and Its Effect Upon Airborne Antenna Noise Temperatures at UHF

Results of an experimental program to determine the characteristics of city noise and its effect on airborne UHF antennas are presented. Coherent sources of energy from communications equipment, radar, navigational aids, etc. are not included in the analysis. Based upon the experimental data, a model of the industrial noise is presented in which an industrial area is considered as a uniformly distributed source of independent radiators. The magnitude of the power density distribution was computed to be 3×10-18 to 1×10-18 W/m2/Hz over the UHF band for all the East Coast cities measured, with the exception of New York City which was 5 to 6 dB higher. Relations have been derived and curves plotted to compute antenna noise temperature increase due to city noise, based upon the metropolitan area and its range from the aircraft.     

Radiative and hybrid cooling of infrared space telescopes

The designs of cold space telescopes, cryogenic and radiatively cooled, are similar in most elements and both benefit from orbits distant from the Earth. In particular such orbits allow the anti-sunward side of radiatively-cooled spacecraft to be used to provide large cooling radiators for the individual radiation shields. Designs incorporating these features have predictedT _tel near 20 K. The attainability of such temperatures is supported by limited practical experience (IRAS, COBE). Supplementary cooling systems (cryogens, mechanical coolers) can be advantageously combined with radiative cooling in hybrid designs to provide robustness against deterioration and yet lower temperatures for detectors, instruments, and even the whole telescope. The possibility of such major additional gains is illustrated by the Very Cold Telescope option under study forEdison, which should offerT _tel5 K for a little extra mechanical cooling capacity.     

Radioisotope powered AMTEC systems

Alkali Metal Thermal to Electric Converter (AMTEC) systems are being developed for high performance spacecraft power systems, including small, General Purpose Heat Source (GPHS) powered systems. Several design concepts have been evaluated for the power range from 75 W to 1 kW. The specific power for these concepts has been found to be as high as 18-20 W/kg and 22 kW/m³. The projected area, including radiators, has been as low as 0.4 m²/kW. AMTEC power systems are extremely attractive, relative to other current and projected power systems, because AMTEC offers high power density, low projected area, and low volume. Two AMTEC cell design types have been identified. A single-tube cell is already under development and a multi-tube cell design, to provide additional power system gains, has undergone proof-of-principle testing. Solar powered AMTEC (SAMTEC) systems are also being developed, and numerous terrestrial applications have been identified for which the same basic AMTEC cells being developed for radioisotope systems are also suitable     

The MESSENGER Spacecraft

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft was designed and constructed to withstand the harsh environments associated with achieving and operating in Mercury orbit. The system can be divided into eight subsystems: structures and mechanisms (e.g., the composite core structure, aluminum launch vehicle adapter, and deployables), propulsion (e.g., the state-of-the-art titanium fuel tanks, thruster modules, and associated plumbing), thermal (e.g., the ceramic-cloth sunshade, heaters, and radiators), power (e.g., solar arrays, battery, and controlling electronics), avionics (e.g., the processors, solid-state recorder, and data handling electronics), software (e.g., processor-supported code that performs commanding, data handling, and spacecraft control), guidance and control (e.g., attitude sensors including star cameras and Sun sensors integrated with controllers including reaction wheels), radio frequency telecommunications (e.g., the spacecraft antenna suites and supporting electronics), and payload (e.g., the science instruments and supporting processors). This system architecture went through an extensive (nearly four-year) development and testing effort that provided the team with confidence that all mission goals will be achieved. Larry E. Mosher passed away during the preparation of this paper.