بررسي تجربي اثر زاويه‌ي پس‌گرا بر ميدان جريان يك بال در سرعت زير صوت

EXPERIMENTAL INVESTIGATION OF SWEEP ANGLE EFFECT ON THE FLOW FIELD OF A WING AT SUBSONIC REGIME

به‌منظور بررسي اثر زاويه‌ي پس‌گرا بر توزيع فشار سطح بالايي يك بال مخصوص، آزمايش‌هاي گوناگوني در رژيم جريان زير صوت انجام گرفت. اين آزمايشات در اعداد رينولدز و زواياي حمله‌ي مختلف، بر روي سه نيم‌مدل بال با نسبت منظري يكسان و زواياي پس‌گراي مختلف صورت پذيرفت. نتايج نشان مي‌دهد كه توزيع فشار سطح بالاي بال به‌دليل وجود جريان‌هاي عرضي ناشي از زاويه‌ي پس‌گراي بال به ايجاد روند متفاوتي نسبت به حالت دو‌بعدي مي‌انجامد. محدود بودن بال سبب ايجاد گردابه‌هايي در نوك بال شده كه باعث تقويت اثرات سه‌بعدي جريان روي بال مي‌شود، به‌نحوي كه موقعيت و مقدار افت فشار روي آن در نزديكي نوك بال نسبت به نقاط ديگر بال متفاوت خواهد بود و كارايي ايرفويل در مقاطع نزديك به نوك بال بيشتر از مقاطع مياني دستخوش تغييرات شده‌ و از حالت بهينه خارج شده است. در زواياي پس‌گراي كم، توزيع فشار روي بال نسبت به تغييرات عدد رينولدز حساسيت كم‌تري دارد.

A series of wind tunnel tests were conducted to examine the flow field over swept wings at subsonic speed under various conditions. For this purpose, three models with sweep angles of 23, 33 and 40 degrees were used. To Increase the Reynolds number, all models were employed as semi span wing types and the effects eliminated by installing the models on a flat plate at their roots. The flat plate was installed in the middle of the test section at a distance from the wall. The tests were performed at various Reynolds numbers and at various angles of attack on all semi-span wings, which had the same aspect ratio but different sweep angles. The results show that increasing the sweep angle causes a decrease in the velocity over the wing surface. The pressure distribution on the upper surface of the wing shows some differences when compared to the 2-D case, due to the existence of the cross flow and wing tip vortices. As the wing sweep angle was varied, the differences varied too. The results indicate that the wing tip vortices amplify the 3D effects of the flow. The pressure reduction in the vicinity of the wing tip deteriorates the performance at that section, compared to other sections of the wing. This phenomenon would eventually cause stall to occur at this section (wing tip) sooner than at other sections. The maximum pressure drop near the wing tip remains unchanged with changing the sweep angle for zero angle of attack only. But, for other angles of attack, the situation is different. For the wing with a sweep angle of 40 degrees, the surface pressure of all sections differs significantly and the maximum pressure is much lower too. However, near the wing tip, the differences are not significant. This phenomenon is related to the effect of wing tip vortices that are more dominant at higher sweep angles. The test results in all three models show that the pressure coefficient decreases with increasing Reynolds, and this phenomenon is most prominent for the highest sweep angle of 40 degrees.