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Savaş uçakları harici yük ayrılmasının panel yöntemi yardımıyla sayısal benzeşimi

Computer simulation of an external store separation via a panel method from a military aircraft

  1. Tez No: 14340
  2. Yazar: ŞEVKİ METO
  3. Danışmanlar: DOÇ.DR. C. RUHİ KAYKAYOĞLU
  4. Tez Türü: Yüksek Lisans
  5. Konular: Uçak Mühendisliği, Aircraft Engineering
  6. Anahtar Kelimeler: Belirtilmemiş.
  7. Yıl: 1991
  8. Dil: Türkçe
  9. Üniversite: İstanbul Teknik Üniversitesi
  10. Enstitü: Fen Bilimleri Enstitüsü
  11. Ana Bilim Dalı: Belirtilmemiş.
  12. Bilim Dalı: Belirtilmemiş.
  13. Sayfa Sayısı: 57

Özet

ÖZET Bu tez konusunu oluşturan askeri uçaklardan harici yük ayrılmasının bilgisayar yardımıyla benzeşimi, 3 aşa mada yerine getirilmiştir. Birinci aşamada harici yükün tahmini ayrılma yörüngesini kapsayacak yeterli büyüklük te seçilen üç boyutlu bölgedeki hız kontrol noktaların daki uçak perturbasyon hızları elde edilmiştir. Uçak perturbasyon hızlarını elde etmek için lineer- leştirilmiş potansiyel denklem ( ( l-Mİ)#xx+#yy+#zz=0), düşük dereceli panelleme yapan bir program yardımıyla uçağın gövde ve taşıyıcı yüzeylerinin panel lennıesi ve her bir panel üzerinde dik hızın sıfır olma şartının uy gulanmasıyla çözülerek panel şiddetleri ve hız kontrol noktalarındaki uçak perturbasyon hızları bulunmuştur. Hız kontrol noktalarındaki uçak perturbasyon hızlarının bulunmasından sonra ikinci aşamada, harici yük de uçak için yapıldığı gibi panel lenerek harici yükün panel geometrileri ve aerodinamik etki katsayıları bulunmuştur. Son aşamada taşıma pozisyonunda bulunan harici yüke etkiyen kuvvet ve momentler TRESSP adlı Fortran programı yardımıyla hesaplanıp 6 serbestlik dereceli hareket denklemleri yeterince küçük bir At zaman araliği için Runge-Kutta yardımıyla integre edilerek harici yükün akım alanındaki yeni konumu elde edilmiştir. Bu işlem harici yükün her yeni konumunda tekrarlanarak tam bir ayrılma yörüngesi elde edilinceye kadar tekrar lanmiştir. Uçağın panel lenmesinde 660 gövde paneli ve 136 kanat paneli, harici yükün panel lenmesinde de 144 gövde paneli kullanılmıştır. Sıkışamaz halde (Mach=0.3) ve sıfır hücum açısında uçan uçaktan bırakılan harici yükün yörüngesi At=.02 sn alınarak hesaplanmış ve elde edilen sonuçların litera türdeki sonuçlarla uyum içinde olduğu görülmüştür. -V±-

Özet (Çeviri)

SUMMARY COMPUTER SIMULATION OF AN EXTERNAL STORE SEPARATION VIA A PANEL METHOD FROM A MILITARY AIRCRAFT The unsteady prediction of store separation from military aircraft is very important because of vital importance in aircraft design. Prior to the I960' s, there were virtually no widely used or generally accepted methods available for pre-f light prediction of store separation trajectories other than wind tunnel testing techniques. With the advent of modern high-speed attack or fighter-bomber jet powered aircraft, the requirement to carry more and more stores, and to release them at higher and higher speeds emerged. High transonic, and even supersonic release speeds became commonplace requirements [1]. Today, there are literally dozens of wind tunnel, wind tunnel/analyses, and purely theoretical analytic methods of store prediction available to the separation analyst or engineer. Each method has its advantages and disadvantages. The purpose of this thesis 'is to design a computer code to predict store separation motion numerically. The computer code, TRESSP, is a six-degree of freedom store trajectory computer code. It uses all required input data such as ejection force, flowfield, store mass properties, and aircraft flight conditions. The complex aircraft flowfield data and aerodynamic coefficients of the store are obtained by using the low order panel code called TRAERO [2]. The panel method is based on linear potential flow theory along the following lines. The surface of the aircraft is subdivided into a large number of panel, each of which contains an aerodynamic singularity dist ribution. Constant source distributions are used on the panels of body components. Double distributions are used on the panels of lifting surfaces to represent lift effects due to incidence and camber. The thickness of lifting surfaces is represented by the source distribu tion whose strengths are constant and a function of the local wing thickness slope. -Vii-The resultant perturbation velocity component w i at grid point i generated by the aircraft is then given by summation of the products of the velocity influence coefficients and the strengths of the aircraft singularities j, i.e. Mx _ a Wt=^Iu;'|i^i (1) >< Here Na denotes the total number of aircraft singulari ties, i.e. the body source and wing vortex singularities and the wing source singularities (whose strengths tf are constant and a function of the local wing thickness slope only). The aircraft perturbation flowfield velocities at the store control points are obtained by interpolation from the velocity components at the grid points and then are incorporated into the boundary condition of tangential flow at the store control points. These store boundary conditions also include the motion of the store. This results in a system of linear equations for the store asijfoj = -(rbooî-Mijpi + nstî-f rij-fi) (2). ' i=l,2,...,Ns Here ns«i is the normal component of the freestream velocity vector at store control point i.nspi is the normal component at store control point i of the aircraft perturbation flow field def ined at the flow grid.nsti is the constant normal velocity induced at store control point by the store wing source distributions (constant and known strengths), and nsfi is the normal component of the velocity at store control point i due to the trans latory and angular motion of the store relative to the aircraft. Ns denotes the total number of the store control points. The right-hand side members of Eq.(2) are known at each store position in the flow grid. In addition, the aerodynamic influence coefficients asij are constant and need not be recalculated at each store position. The unknown body source and wing vortex singularity strengths Vs] of the store are obtained from the solution of Eq. (2) then the store loads are calculated in the usual manner and applied in the equations of motion to obtain the separation characteristics. The equations consist of the following three force equations : -iri a-w s w g w g u* w“ =¦ Fx = Fy = Fz (3) and Euler' s equations IllWl”- (122 - l33)W2W3= Ml l22W2°- (133 - Ill)W3Wl= M2 l33W3°- (111 - I22)W1W2= M3 (4) The forces Fx, Fy, and Fz include aerodynamic, gravita tional, and any other external forces. The moments Ml, M2, and M3 are the moments of these forces about the store referance axes. (Fig. 1). In, I22, and I33 are the moments of inertia about x, y.and z respectively. And wi, W2, and W3 are the angular velocity about the referance axes. Figure 1. Store referance axes. -ix-A control point is associated with each panel. The three perturbation velocity components, induced by each of the singularities, are calculated at the control points, and from these the components normal to the panels are formed. The magnitude of the normal velocity component induced in control point i by the singularity j of unit strength is called the aerodynamic influence coefficient aij. The resultant normal velocity component at control point is then given by summation of the products of the influence coefficients aij and the strengths of the singularities jk In these way a system of linear equations with coefficients aij is obtained, relating the magnitude of the normal veloci ties at the control points to the unknown singularity strengths. The singularity strengths that satisfy the boundary condition of tangential flow at the control points for a given Mach number and angle of attack are determined by solving this system of equations using an iterative procedure. In order to treat external stores of arbitrary geometry, the stores are represented by singularity panels, in the same manner as the aircraft- Computer run time per store trajectory is reduced drastically by employing a flow grid method in combination with approximations regarding the aircraft-store interaction. In order to treat the aircraft-store interaction, TRESSP was fitted with the“flow grid option.”In this option, the isolated store moveö through the nonuniform flowfield consisting of the freestream plus the pertur bation field created by the aircraft. The nonuniform flowfield is defined at the mesh points of a three di mensional orthogonal grid covering the separation region of interest. The grid is aligned with the aircraft co ordinate system. The aircraft singularity strengths are determined for the aircraft in isolation and remain constant. This is equivalent to assuming that the store is sufficiently small with respect to the aircraft that its influence on the aircraft can be neglected. Let wij be any of the velocity component (u,v,w) induced at grid point i by aircraft singularity j of unit strength; wij is called a velocity influence coefficient. These velocity influence coefficients at the grid points need to be calculated only once and then can be used for any subsequent store position and orientation and for any angle of attack of the aircraft under the condition that the freestream Mach number is constant. -x-Store Separation The separation characteristics of the released store are obtained as follows. First, TRESSF calculates the forces and moments acting on the store in the release position. The program numerically integrates the six-degree-of-freedom equations of motion for the store for a specified small time interval At by using Runge-Kutta method of fourth step type to arrive at a new store position and attitude in the flow grid. Then the loads acting on the store at its new position and attitude are computed and the equation of motion again solved for a small time interval [3]. The aircraft perturbation flowfield was obtained with 660 body panels and 136 lifting surface panels. The external store loads and trajectory calculations were performed with 144 store body panels. Store trajectories have been obtained using a time step At=0.02 s. Initially (at time t=0), the store longitudinal axes is paralel to the longitudinal axis of aircraft fuselage. The initial values of the angular velocity components of the store are zero. The mass of the store is 50 kg* and the center of mass is located at the store midpoint. The aircraft is in level flight (zero climb or dive angle) at an altitude of 2000 m. -xi-

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