航空学报
航空學報
항공학보
ACTA AERONAUTICA ET ASTRONAUTICA SINICA
2009年
12期
2288-2294
,共7页
航空航天推进系统%吸气式高超声速飞行器%二元进气道%风洞试验%数值模拟
航空航天推進繫統%吸氣式高超聲速飛行器%二元進氣道%風洞試驗%數值模擬
항공항천추진계통%흡기식고초성속비행기%이원진기도%풍동시험%수치모의
aerospace propulsion system%airbreathing hypersonic vehicle%two-dimensional inlet%wind tunneltest%numerical simulation
针对飞行马赫数为6.00的二元进气道模型开展了高焓脉冲风洞试验研究,分析了进气道在不设置反压和设置反压两种情况下的激波结构、内通道皮托压分布及隔离段出口的性能,并结合数值仿真分析了通道内的流场特性.研究结果表明:在无反压情况下,进气道内通道激波反射明显,靠近下壁面的皮托压值均低于其他测点,在隔离段出口截面,靠近侧壁皮托压有所降低;在一定反压条件下,结尾激波系上传至隔离段内,结尾激波位置不对称;堵塞度为62%的反压条件下,结尾激波系位于喉道位置,隔离段出口截面下半部分已经是亚声速流动;在来流马赫数Ma=6.07,迎角α=4.5°无反压情况下,隔离段出口总压恢复系数为0.477,平均马赫数为2.72,增压比为44,流量系数为0.81,表明进气道性能良好.
針對飛行馬赫數為6.00的二元進氣道模型開展瞭高焓脈遲風洞試驗研究,分析瞭進氣道在不設置反壓和設置反壓兩種情況下的激波結構、內通道皮託壓分佈及隔離段齣口的性能,併結閤數值倣真分析瞭通道內的流場特性.研究結果錶明:在無反壓情況下,進氣道內通道激波反射明顯,靠近下壁麵的皮託壓值均低于其他測點,在隔離段齣口截麵,靠近側壁皮託壓有所降低;在一定反壓條件下,結尾激波繫上傳至隔離段內,結尾激波位置不對稱;堵塞度為62%的反壓條件下,結尾激波繫位于喉道位置,隔離段齣口截麵下半部分已經是亞聲速流動;在來流馬赫數Ma=6.07,迎角α=4.5°無反壓情況下,隔離段齣口總壓恢複繫數為0.477,平均馬赫數為2.72,增壓比為44,流量繫數為0.81,錶明進氣道性能良好.
침대비행마혁수위6.00적이원진기도모형개전료고함맥충풍동시험연구,분석료진기도재불설치반압화설치반압량충정황하적격파결구、내통도피탁압분포급격리단출구적성능,병결합수치방진분석료통도내적류장특성.연구결과표명:재무반압정황하,진기도내통도격파반사명현,고근하벽면적피탁압치균저우기타측점,재격리단출구절면,고근측벽피탁압유소강저;재일정반압조건하,결미격파계상전지격리단내,결미격파위치불대칭;도새도위62%적반압조건하,결미격파계위우후도위치,격리단출구절면하반부분이경시아성속류동;재래류마혁수Ma=6.07,영각α=4.5°무반압정황하,격리단출구총압회복계수위0.477,평균마혁수위2.72,증압비위44,류량계수위0.81,표명진기도성능량호.
The performance of a fixed-geometry two-dimensional mixed-compression hypersonic inlet is investigated at Mach number 6.00 in the high-enthalpy impulse wind tunnel. The investigation focuses on the structure of the shocks, the distribution of Pitot-pressure ratio at the duct and the performance of the inlet at the exit of the isolator. Characteristics of the internal flow are analyzed in combination with numerical simulation. Results indicate: (1) With no pressure imposed by the mass flow meter, there is a distinct reflected shock structure in the isolator, and the Pitot-pressure of the probe at the exit section of the isolator near the ramp side is lower than that of all other probes. Meanshile, the Pitot-pressure of the rake near the sidewall is lower than others. (2) The terminal shock is propagated in the isolator when the back pressure ratio increases to a certain value. The terminal shock is asymmetric in the isolator. When the throttling ratio increases to 62%, the terminal shock is located at the section of the throat and the Pitot-pressure profiles at the exit of the isolator show that the flow speed for at least a half region in the isolator height is already subsonic. (3) At Ma=6. 07,α=4. 50° and with no back pressure, the total pressure recovery coefficient and the mass flow ratio are 0.477 and 0.81 respectively at the exit of the isolator, while the average Mach number is 2.72 and the pressure ratio between the flow at the exit of the isolator and free stream is 44.
