Aerodynamic Analysis of Camber Morphing Airfoils in Transition via Computational Fluid Dynamics

Abstract

In this paper, the authors analyze an important but overlooked area, the aerodynamics of the variable camber morphing wing in transition, where 6% camber changes from 2% to 8% using the two airfoil configurations: NACA2410 and NACA8410. Many morphing works focus on analyzing the aerodynamics of a particular airfoil geometry or already morphed case. The authors mainly address "transitional" or "in-between" aerodynamics to understand the semantics of morphing in-flight and explore the linearity in the relationship when the camber rate is gradually changed. In general, morphing technologies are considered a new paradigm for next-generation aircraft designs with highly agile flight and control and a multidisciplinary optimal design process that enables aircraft to perform substantially better than current ones. Morphing aircraft adjust wing shapes conformally, promoting an enlarged flight envelope, enhanced performance, and higher energy sustainability. Whereas the recent advancement in manufacturing and material processing, composite and Smart materials has enabled the implementation of morphing wings, designing a morphing wing aircraft is more challenging than modern aircraft in terms of reliable numerical modeling and aerodynamic analysis. Hence, it is interesting to investigate modeling the transitional aerodynamics of morphing airfoils using a numerical analysis such as computational fluid dynamics. The result shows that the SST k-ω model with transition/curvature correction computes a reasonably accurate value than an analytical solution. Additionally, the CL is less sensitive to transition near the leading edge in airfoils. Therefore, as the camber rate changes or gradually increases, the aerodynamic behavior correspondingly changes linearly.

Keywords: CFD; airfoil; analytical and numerical; benchmark; camber morphing; computational fluid dynamics; transition.

Figure 1

Cross-sectional geometry and description of an example airfoil [2].

Figure 2

NACA2410-2410 wing and its dimensions.

Figure 3

NACA8410 (at the wing tip) to NACA2410 (at the fuselage) wing under stress model.

Figure 4

Mesh Independence Study (NACA Airfoil, Re = 6 Million,$\mathsf{\alpha }$ = 0°).

Figure 5

(a) input face and (b) outlet face. (c) the left-side symmetry faces and (d) the right-side symmetry face. (e) Airfoil inside the computational domain.

Figure 6

(a) $C_L$ and (b) $C_D$ vs. varying the AoA.

Figure 7

Pressure contour of NACA2410 with AoA (a) 1° and (b) 8°.

Figure 8

Velocity contour of NACA2410 with AoA (a) 1° and (b) 8°.

Figure 9

Turbulence intensity contour of NACA2410 with AoA (a) 1° and (b) 8°.

Figure 10

Pressure distribution around NACA2410 with varying AoA while x ticks from 0 with increment of 0.32; (0, 0.32, 0.64, 0.96, 1.28, 1.6, and 1.92, respectively).

Figure 11

$C_D$ (b) and $C_L$ (a) of NACA2410 and NACA8410 vs. varying AoA.

Figure 12

Pressure contour of NACA8410 with AoA (a) 1° and (b) 8°.

Figure 12

Pressure contour of NACA8410 with AoA (a) 1° and (b) 8°.

Figure 13

Velocity contour of NACA8410 with AoA (a) 1° and (b) 8°.

Figure 13

Velocity contour of NACA8410 with AoA (a) 1° and (b) 8°.

Figure 14

Turbulent kinetic energy, $k$ contour of NACA8410 simulation with AoA (a) 1° and (b) 8°.

Figure 14

Turbulent kinetic energy, $k$ contour of NACA8410 simulation with AoA (a) 1° and (b) 8°.

Figure 15

Pressure distribution of the NACA8410 with varying AoA while x ticks from 0 with increment of 0.32; (0, 0.32, 0.64, 0.96, 1.28, 1.6, and 1.92 respectively).

Figure 16

Input (deformation) distribution to morph NACA2410 to NACA8410 along the wing.

Figure 17

Output (maximum shear stress) distribution to morph NACA2410 to NACA8410 along the wing.

Figure 18

Sectional output (maximum shear stress) distribution to morph NACA2410 to NACA8410 along the wing.

Figure 19

(a) Y axial output (deformation) distribution and (b) Z axial output (deformation) distribution) to morph NACA2410 to NACA8410 along the wing.

Figure 20

(a) $C_L$, (b) $C_D$ of NACA2410, NACA8410, & NACA2410 to NACA8410 transition vs. AoA.

Figure 21

Pressure contour of the NACA2410 to NACA8410 transition at AoA 5°.

Figure 22

Velocity contour of NACA2410 to NACA8410 transition at AoA 5°.

Figure 23

Turbulent kinetic energy, $k$ contour of NACA2410 to NACA8410 transition at 5° AoA.