高進氣道 的英文怎麼說
中文拼音 [gāojìnqìdào]
高進氣道
英文
high intake-
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， 2005The effect of exit back pressure and length / height ratio of isolator on forebody / inlet flow field is analyzed and discussed in the paper
本文還研究了隔離段出口反壓及隔離段長高比對進氣道流場的影響。The design point of the designed forebody / inlet is flight mach number m = 6 and flight attitude h = 25km. the three - ramp, mixed external and internal compression hypersonic / inlet is chosen as representative geometry to be designed. nine types of hypersonic / inlet model with difference forebody length and total turning angle are obtained that the incidence angle of the first ramp and height of isolator are given
根據飛行任務要求，以飛行馬赫數6 ，飛行高度25km為設計點，給定前體預壓縮角、隔離段高度，對具有不同前體長度和不同總壓縮轉角的帶隔離段的前體進氣道進行了設計。With account for the effect of temperature and shock - boundary layer interaction, the forebody / inlet with plane lip and wedge lip are designed by using the design methods of constant shock wave intensity and constant shock wave angle. in the design, the incidence angle of the first ramp, height of isolator, length of forebody and total turning angle of the designed forebody / inlet models are given
另外，根據優選結果，給定前體長度、前體預壓縮楔角、總壓縮轉角及隔離段高度，分別用等激波強度和等激波角度的方法設計了唇口平直和唇口帶楔角的前體進氣道。The calculation results indicate that the starting performance of forebody / inlet with wedge - shaped lip is better than that with plane lip at low flight mach number. at high flight mach number, the performance of the forebody / inlets designed by the two methods is approximate. at low flight mach number, the starting performance of forebody / inlet designed by the method of constant shock wave intensity is better
結果表明：唇口帶楔角的進氣道在低飛行馬赫數下的起動性能優于唇口平直的進氣道；採用等激波強度和等激波角度方法設計的進氣道在高飛行馬赫數下的性能相近，在低飛行馬赫數下，採用等激波強度方法設計的進氣道起動性能較優。In addition to a longer nose, aerodynamic modifications feature kfir - style small nose side - strakes to prevent yaw departure at high aoa, a pair of fixed delta canards on the upper parts of the air intakes, dog - tooth outboard leading - edge extensions, and short fences replacing leading - edge slots
除了一個加長的機鼻外，氣動布局的修改包括：在機鼻兩則裝上可以防止在高攻角下脫離偏航的「幼獅」式小邊條，一對固定在進氣道的三角鴨翼，鋸齒型外翼前緣，和代替前緣翼槽的短翼刀。Cross - section design of a controllable hypersonic inlet. the research is being done at the nanjing university of aeronautics and astronautics
南京航空航天大學正在進行一個可控高超音速進氣道斷面設計研究。Finally, the inlet port and combustion chamber designed and machined is tested on 4100qb. the paper mostly investigates the change of performance targets of the engine that three different grooved chambers including two different re - entrant model and one w model match with various intake port and nozzles, in order to realize the influence of chamber geometry on combustion process. the results analyzed have indicated that re - entrant grooved chambers can enhance the turbulent intension in the chamber, improve the combustion proces of diesel engine on the altiplano, being groove, re - entrant and convex, and have the farther developing potential
著重研究兩種縮口、一種直口矩形環槽燃燒室與不同進氣道、噴油嘴的組合方案發動機的動力性、經濟性及排放指標的變化，確定燃燒室的幾何形狀對燃燒過程的影響。分析結果表明，由於有矩形環槽、燃燒室縮口以及燃燒室底部凸臺，縮口矩形環槽燃燒室中的紊流得以加強，改善了高原柴油機的燃燒過程，這種燃燒室有進一步發展的潛力。Sum up, the study of intake port is significant for the design of intake port and the development of engine. at present, the study of flow characters in intake port of engine mainly depends on experiment. numerical simulation is not universal, especially for the engine with dual - intake port
目前，國內對發動機進氣道的流動特性研究手段主要依靠實驗，在實驗臺上進行大量的實驗，既費時又費力，實驗方法顯得過于陳舊過時，已不能適應現代高性能發動機研製工作的需要，因此，採用多維數值模擬技術是一種有效的研究手段。Based on the variables of inlet length, compression ratio, inlet opening phase and the exhaust closing phase, multi - objective optimization was conducted for power, torque, combustion noise and fuel consumption indicators with constraint conditions of fuel consumption rate and the maximum breakout pressure
以進氣道長度、壓縮比、進氣開啟相位、排氣關閉相位為可變因素；以比油耗、最高爆發壓力為約束條件，針對滿足功率、扭矩、燃燒噪聲和油耗指標進行多目標優化設計。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。A ceramic coating is applied to the inner surface of the inlet duct to protect the composite from the high temperature
由於復合材料耐高溫性能差，該進氣道在內部使用防熱陶瓷塗層進行熱防護。While the speed is very large, the fuel - air mixing behind the airstream is decreased and the span that the fuel flow in the second combustor is reduced, so the combustion efficiency is decreased, h ) if the fuel streams impinge with the airstreams directly at the air - inlet exit, it make against increasing the combustion efficiency, i ) increasing air - to - fuel ratio within proper range can increase the combustion efficiency
增加燃氣噴射速度有利於增強迴流區強度，頭部燃燒溫度上升，但速度太大則會減弱燃氣同空氣在進氣道下游的摻混燃燒，減少燃氣在室內停留時間，燃燒效率降低； 8燃氣射流與空氣流在進氣道出口直接撞擊不利於燃燒效率的提高； 9在適當范圍內增大空燃比能顯著提高燃燒效率。The mixing of air and fuel can be improved, d ) adding air distributary valve at exit of air - intake can increase the amount of air bifurcated into the dome region, so the value of combustion efficiency is increased, e ) increasing gas generator nozzle number can improve combustion characteristics in the dome region and produce good condition for combustion progress, f ) when the fuel streams are ejected into the second combustor with a angle, the combustion efficiency is increased, g ) when the speed of fuel streams increased, the intension and temperature of recirculation region is improved
兩個進氣道在補燃室周向成180度布置，有利於增強燃氣同空氣的摻混； 4在進氣道出口增加空氣分流擋板，有利於增加進入頭部的空氣流量，提高燃燒效率； 5增加燃氣噴管數量能增強燃氣同空氣在頭部的燃燒效果，為燃燒的進行創造良好的條件； 6It is no effect on a sidewall compression inlet ' s starting characteristics while a starting sidewall compression inlet is moved cowl to increase interior contraction ratio appropriately. back pressure effect on characteristics of self - starting for hypersonic sidewall compression inlet is tested in mach 3. 85 wind tunnel. the hypersonic sidewall compression inlet with interior contraction ratio 1. 24 can self - start. unstart hypersonic sidewall compression inlet caused by high back pressure can self - start by moving cowl to decrease interior contraction ratio
最後，實驗研究了反壓對側壓式進氣道自起動特性的影響，內收縮比1 . 24的側壓式進氣道可以實現自起動並且實驗驗證了由於反壓過高而不起動的側壓式進氣道，通過移動唇口板以減小內收縮比可以實現側壓式進氣道的自起動。Back pressure propagation mode and maximum working back pressure of hypersonic inlet isolator
高超聲速進氣道隔離段反壓的前傳模式及最大工作反壓Furthermore, there are complex structures such as the engine fans at the end of the inlets. it ' s hard to deal with such problems if we use high frequency method or low frequency numerical value method only
對于厘米波段雷達而言，進氣道屬于電大尺寸目標，其終端有渦輪葉片之類的復雜結構，單用高頻近似法或低頻數值法都難以解決。The improvement of the flow character in intake port improves the dynamic characters and economy. all this can be used to reduce the cost
提高進氣道的流動性能，可以提高發動機的動力性和經濟性，這是降低經濟成本的有效方法之一。The results show the exit back pressure and length / height ratio of isolator have great impact on exit flow field and starting performance of inlet. high back pressure of inlet will result in unstart
研究表明：隔離段長高比對進氣道出口參數影響較大；隔離段出口反壓不能太高，否則會引起進氣道不起動。Scramjet combustor may realize the subsonic - combustion mode and supersonic - combustion mode respectively based on the control of fuel in variable flight conditions. isolator has great effect on mode transition and preventing unstart of inlet. inlet can be unstarted by raising the back pressure result from chemical energy release in the combustor
計算研究表明，在相同的來流條件下，供油規律對燃燒室的工作模態有著重大的影響；在不同的來流條件下，可以通過調節供油規律使燃燒室分別實現亞燃模態和超燃模態；隔離段對防止燃燒引起的壓力升高對進氣道的干擾和燃燒室的模態轉換具有重要的作用。分享友人