高超聲速的 的英文怎麼說

中文拼音 [gāochāoshēngsùde]
高超聲速的 英文
hypersonic
  • 高 : Ⅰ形容詞1 （從下向上距離大; 離地面遠） tall; high 2 （在一般標準或平均程度之上; 等級在上的） above...
  • 超 : Ⅰ動詞1 （越過; 高出） exceed; surpass; overtake 2 （在某個范圍以外; 不受限制） transcend; go beyo...
  • 速 : Ⅰ形容詞（迅速; 快） fast; rapid; quick; speedy Ⅱ名詞1 （速度） speed; velocity 2 （姓氏） a surna...
  • 的 : 4次方是 The fourth power of 2 is direction
  • 高超 : excellent; exquisite; superb
  1. Tahir r b, molder s and timofeev e v. unsteady starting of high number air inlets - a cfd study [ r ]. aiaa 2003 - 5191

    梁德旺，袁化成，張曉嘉.影響高超聲速進氣道起動能力的因素[ c ] / / 2005年沖壓發動機技術交流會論文集， 2005
  2. At last, the effect on the outer ballistic trajectory characteristic after the ablation of hypersonic projectile wing has been analyzed

    最後，初步分析了高超聲速彈翼表面發生燒蝕后對其外彈道特性的影響。
  3. Numerical simulation of 3d hypersonic thermochemical nonequilibrium flow

    高超聲速三維熱化學非平衡流場的數值模擬
  4. The paper provides an description of stage separaton of hypersonic vechicle

    本文簡述了高超聲速飛行器飛行過程中的分離問題。
  5. In this paper, some flows around naca 0012 aerofoil, m6 wing and two different blunt vehicles are simulated, the results show that the present algorithm possesses good characteristics and capability for normal mach number and for hypersonic of simple configurations

    通過對naca0012翼型、 m6機翼、鈍頭以及雙鈍錐?柱體的算例驗證，本文能很好的解決一般馬赫數問題，能夠很好的求解簡單外形的高超聲速問題，對于復雜外形飛行器還需要進一步進行研究驗證。
  6. In this thesis, based on pershing ii surface to surface missile, a new kind of ballistic missile was designed with an additional rocket engine, which can be ignited twice. firstly, aerodynamic computational models of missile body and warhead which reentry with supersonic are built according to the task requirements ; secondly, the propulsion system model of missile is built whose first two stages are solid rocket engines and the third stage is liquid - solid combined rocket engine. the nozzle and the shape of the engine are designed to meet the needs of the populsion project ; thirdly, the trajectory model of the mass point is built and a wavy trajectory is designed & optimized ; finally, the ability of a missile ' s breaking through defence is analyzed

    以美國潘興導彈為原型，增加可兩次點火的末級發動機，改裝成具有跳躍能力的地地彈道導彈；首先，根據任務需求，建立了導彈的氣動模型，並建立了彈頭再入時高超聲速氣動模型；其次，建立了導彈推進系統模型，前兩級採用了固體火箭發動機，第三級採用了固?液組合火箭發動機，並在總體方案要求下，對發動機噴管和外形進行了設計；第三部分，建立了導彈質點彈道模型，設計了一條跳躍式彈道，並對跳躍式彈道進行了優化設計；最後，對導彈進行了突防能力分析，從分析的結果可以看出，跳躍式彈道的突防能力比常規的拋物線彈道要強。
  7. Hypersonic propulsion test facility is important for the scramjet ground tests

    摘要高超聲速推進風洞是進行超燃沖壓發動機模型地面模擬的重要試驗設備。
  8. Scramjet is the perfect propulsive device to hypersonic vehicle, and is a hot issue in current hypersonic technology area

    超燃沖壓發動機是高超聲速飛行器理想的動力裝置，是當前高超聲速技術領域的研究熱點。
  9. In the range of flight mach number mo = 3 ~ 7, in order to take the advantage of ramjet and scramjet, a fix - shaped dual - mode scramjet, in which both subsonic and supersonic combustion can be obtained, is used for hypersonic vehicle

    在飛行馬赫數m _ = 3 - 7的范圍內，為充分發揮沖壓發動機的性能優勢，高超聲速飛行器採用了既能在亞聲速燃燒模態，又能在超聲速燃燒模態下工作的固定幾何形狀的雙模態沖壓發動機。
  10. Both advantages of rocket engine and air - breathing engine are integrated into one kind of propulsion system, called rocket - based - combined - cycle ( rbcc ), which has multiple operating woke modes, including ejecting, subsonic combustion, supersonic combustion and rocket mode, with each mode operating at varying flying stage. rbcc has prominent potential of high efficiency and low cost, and so it is recognized as an advanced propulsion system of future single - stage - to - orbit reusable spacecraft and hipper - sonic missile. it has been broadly investigated in foreign countries and has been progressed into small and full - scale flying stage

    火箭基組合動力循環（ rbcc ， rocketbasedcombinedcycle ）將傳統的火箭發動機和吸氣式發動機的優點集中組合到一個具有多種工作模態（包括引射、亞燃、超燃及火箭模態）的發動機里，在不同的飛行階段啟用不同的工作模態，因此具有大幅提高航天推進系統經濟性與高效性的潛在優勢，並可能發展成為下一代單級可重復使用航天器以及高超聲速導彈武器的動力系統。
  11. The tests were conducted in the hypersonic low density wind tunnel at nominal test conditions of mach 16, stagnation temperature 923k, stagnation pressure 1. 40mpa and 7. 30mpa. heat - transfer data were obtained on a hemisphere model, a sharp cone and a big blunt cone respectively by means of infrared thermal mapping techniques, that of a 0. 5mm thickness blunt cone by virtues of thermocouples. furth ermore, heat - transfer on all those models was calculated with the theoretical method

    最後在名義m _ = 16 、 t _ 0 = 923k 、 p _ 0 = 1 . 40mpa及7 . 30mpa的高超聲速低密度風洞中，利用紅外熱圖技術獲得了半球圓柱、尖錐、大鈍頭三個模型表面熱流分佈，利用薄壁法技術得到了一壁厚為0 . 5mm的鈍錐模型表面的熱流分佈，並通過工程理論方法計算了模型表面的氣動熱，把理論計算結果與上述試驗結果比較，幾者符合得較好。
  12. Although a dual - mode scramjet ' s configuration is simple and mainly consists of inlet, combustor and wake nozzle, its working process is complicated, especially in the combustor, involving a lot of subjects, including hypersonic aerodynamics, combustion chemistry, etc. the inner flow of a combustor is three - dimensional and complicated, including the interaction of shock wave, deflagration, vortex and boundary layer, and so on

    它涉及到高超聲速空氣動力學、燃燒化學、擴散傳質等多門學科；其內部的實際流動是復雜的三維流動過程，充滿著激波、膨脹波、燃燒波、各種渦系、附面層及其相互之間的干擾，因此，燃燒室問題是整個發動機研究的關鍵所在。
  13. In the condition of non - axisymmetrical stage separation in hypersonic flight, the problems of dynamics and kinematics of the validating vechicle and the launch vechicle are primarily analysised. it includes the movement of the center of mass and the rigid body of the vechicles, the relative movement of the different vechicles, the relative movement of the special parts of the different vechicles, the load analysis and the probable collision in the process of stage separation

    研究了在高超聲速狀態下非對稱級間分離過程中，驗證飛行器與運載火箭分離時的相關運動學和動力學問題。包括:飛行體質心運動和剛體運動;不同飛行體之間的相對運動;不同飛行體上特殊部位之間的相對運動;分離過程中各相關飛行體上的載荷分析和分離過程中可能發生的碰撞問題。
  14. The flowfields around the reentry blunt body are simulated numerically by solving the navier - stokes with the various thermal and chemical model, such as perfect gas model, vibrational excitation model, equilibrium gas model, and the non - equilibrium gas model of one temperature, two temperature and three temperature

    本文採用完全氣體模型、振動激發氣體模型、平衡氣體模型、一溫度非平衡氣體模型、兩溫度非平衡氣體模型和三溫度非平衡氣體模型進行了鈍體高超聲速繞流流場的數值計算。
  15. A hypersonic pulse tunnel is designed in this paper, which has multiple running modes based on gun tunnel and shock tube theory

    本文設計了一座具有多種運行方式的高超聲速大型脈沖風洞，該風洞可以採用炮風洞和激波管兩種運行方式。
  16. The code calibration is completed by comparing the numerical results with data from two experiments, one of which is test of a 2 - d hypersonic inlet with internal compression and the other a sidewall compression inlet. good agreement of numerical and experimental results proves the possibility and credibility of the numerical method

    利用國外文獻公布的二維進氣道實驗結果和本課題組側壓式進氣道炮風洞實驗獲得的實驗數據對本文的計算方法進行了校驗，證明了本文所採用的方法應用於高超聲速進氣道計算的可行性和可靠性。
  17. On the mathematical simulation of non - equilibrium hypersonic flow around the flying vehicle

    關于飛行器高超聲速不平衡氣體繞流的數值模擬
  18. The application of uniform experimental design to configuration design of hypersonic aircraft

    均勻試驗設計法在高超聲速飛行器外形設計中的應用
  19. The model is solved with the finite difference method, the simulation software is organized to simulate the transient thermal response of the front region of projectile wing such as those launched by electrothermal chemical guns

    採用有限差分法進行數值求解，並編制了較為通用的模擬軟體，用於模擬諸如由電熱化學炮發射的高超聲速穿甲彈彈翼前緣的瞬時熱效應。
  20. To prove the accuracy of the mach number, and the parameter homogeneity of the design nozzle " s exit, cfd calculate has carried on the design results. under the condition of supersonic and hypersonic flow, and a certain range of temperature, and mach number, the conclusion of the influence of specific heat to nozzle design is drawn

    為了驗證所設計的噴管出口馬赫數的大小和噴管出口流場的均勻性，採用nnd格式和b l湍流模型求解雷諾平均n - s方程，對設計結果進行了cfd驗算，得出了在一定溫度范圍內，超音速、高超聲速流動的條件下，不同馬赫數范圍內變比熱容對噴管型面和噴管出口馬赫數的影響。
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