الفهرس | Only 14 pages are availabe for public view |
Abstract Traditionally, early in the design studies, when many concepts of aircraft configurations are being considered, designers use their experience and historical data to include their considerations in the concept. However, this approach is limited to more conventional configurations and can be very time consuming for this stage of the design process. Once the specialists get involved, more detailed numerical methodologies are used. However, those methods cannot yet respond to the “dozen a day” type configuration evaluations desired in the initial conceptual design stages. The aim of this work is to develop an easy to use and low time consuming solver for predicting the aerodynamic characteristics and estimating the stability and control derivatives of a complete configuration aircraft in subsonic flow regime. The developed solver is integrated with pre- and post- processing modules to prepare the geometric data of the computational model and visualize the results respectively. This solver can be used easily by designers for fast evaluation of any aircraft configuration in the initial conceptual design stages. The present solver computes the numerical solution for the integral form of Laplace’s equation using the vortex lattice method by employing vortex ring elements. The Boit-Savart law is used to compute the induced velocity and the Kutta- Jukowski theorem is applied to compute the force at each vortex ring element. The boundary conditions are that there is no normal velocity component over the surface and the aircrafts operate in subsonic flow at low angles of attack flight regime In order to validate the developed potential flow solver, three different wing planforms were used as the test cases. First, a rectangular wing of AR = 2, b = 12 m, and CMAC = 6 m at α = 5 and 7 degrees. Second, a swept back wing of TR = 1, SA = 450, AR =5, b = 10 m, CMAC = 2 m, and S = 20 m2 at angles of attack in the range from 0 to 10 degrees. Finally, a swept tapered wing of TR = 0.5, SA = 450, AR = 2, α = 5.00, b = 12 m, CMAC = 6 m, and S = 72 m2. All the results were compared with the published experimental and computational results. The influences of varying the different parameters on the computed force and moment coefficients, and the load distribution over the wing span were studied. The effects of the planform mish divisions were investigated by performing different computations on a swept back wing using chordwise divisions in the range from 2 to 16 and spanwise divisions in the range from 2 to 16. The wing geometric parameters were studied by testing different wing configurations of aspect ratio (AR) in the range from 6 to 12, taper ratio (TR) in the range from 0 to 1, swept angle (SA) in the range from 0 to 60 degrees, dihedral angle (DIH) in the range from 10 to 40 degrees, angle of attack (AOA) in the range from 6 to 12 degrees, and ground effect at different values of the ratio between the height above the ground and the chord (h/c) in the range from 0.2 to 2. The research was conducted to estimate the values of the stability and control derivatives for the three-dimensional model of multi-element lifting surface with trailing edge control surfaces using the developed solver. The stability and control derivatives were estimated for conventional wing configuration aircrafts like Cessna 172, Boeing 747, and F18; and non-conventional wing configuration aircrafts like boxwings. |