CH. 4 Drag

5. CH. 4 Drag ▣ Drags * surface friction drag (표면마찰항력) * trailing-vortex (induced) drag (꼬리와류항력, 유도항력) * boundary layer normal pressure (form) drag (경계층 수직압력항력, 형성항력) * wave drag (조파항력) ▣ Two basic constituents of drag * force due to the pressure distribution * force due to viscous shearing ▣ Boundary Layer Drag (Profile Drag) = boundary layer normal pressure (form) drag + surface friction drag

8. 1./ Drag Coeefficient • ♣ • where • ♣ • ♣ dependence of on Reynolds number ; weak over a wide range • ♣ dependence of on Mach number ; • 2./ Boundary Layer Normal Pressure (Form) Drag (p94 Fig. 4.1) • ♣ In real case shown in Fig. 4.1(b) and Fig. 4.2, the streamline pattern and pressure distribution are not symmetrical, and a wake of slow-moving air is formed at the rear. • ♣ Even if the flow does not separate, an adverse pressure gradient promotes a rapid degradation of available energy in the

9. boundary layer, resulting in a reduction in pressure over the rear. Thus, on average, there is a lower pressure on the rear of the section than on the front, and therefore, there is now a net drag force, which is known as the boundary layer normal pressure (form) drag. ♣ In the boundary layer and the wake, the speed and the pressure can be simultaneously lower than in the free stream values.

20. 3./ Reducing Boundary Layer Normal Pressure Drag ▣ In order to reduce the boundary layer normal pressure drag, it is important to ensure that the pressure gradient is not strongly adverse, which means that the tail of the body should reduce in depth or cross-sectional area gradually.  streamlining ▣ advantages of streamlining ▣ reducing frontal area * Strongly unfavorable pressure gradients can be avoided by making all parts of the aircraft as thin as possible: in other words, by reducing the frontal area. * There is always an optimum compromise between decreased boundary layer normal pressure drag resulting from reduced frontal area, and increased surface friction drag caused by the increased surface area.

26. Fluid dynamic forces are comprised of pressure and friction effects. • Often useful to decompose, • FD = FD,friction + FD,pressure • CD = CD,friction + CD,pressure • This forms the basis of ship model testing where it is assumed that • CD,pressure = f(Fr) • CD,friction = f(Re)

28. Example: Automobile Drag Scion XB Porsche 911 CD = 1.0, A = 25 ft2, CDA = 25ft2 CD = 0.28, A = 10 ft2, CDA = 2.8ft2 • Drag force FD=1/2V2(CDA) will be ~ 10 times larger for Scion XB • Source is large CD and large projected area • Power consumption P = FDV =1/2V3(CDA) for both scales with V3!

29. Flow is strong function of Re. • Wake narrows for turbulent flow since TBL (turbulent boundary layer) is more resistant to separation due to adverse pressure gradient. • sep,lam ≈ 80º • sep,turb ≈ 140º

32. 4./ Influence of Boundary Layer Type ▣ As a turbulent boundary layer generates a greater surface friction drag than a laminar one, it is once again necessary to strike the correct balance between reduction in form drag, and rise in surface friction drag. ▣ Note that roughening the surface will only have a beneficial effect if the Reynolds number is in the critical region.

37. 5./ Low Drag Wing Sections

39. Airfoil Families 1. NACA Four-Digit Series 1] Numbering System The four-digit airfoil geometry is defined, as the name implies, by four digits ; - 1st digit ; the maximum camber in percent of chord - 2nd digit ; location of the maximum camber in tenth of chord - last two digits ; maximum thickness in percent of chord 2] example NACA 4412 ; 12% thick airfoil having a 4% camber located 0.4c from the leading edge - 1st 4 ; max camber of 4% of chord

40. - 2nd 4 ; max camber's location from LE = 0.4c - 3rd & 4th 12 ; max thickness in % of chord 2. NACA Five-Digit Series 1] Numbering System ♣ 1st digit ; - The 1st digit multiplied by 3/2 gives the design lift coefficient in tenth of the airfoil. - Design lift coefficient in tenth is three-halves of the 1st integer. - the amount of camber in terms of the relative magnitude of the design lift coefficient ♣ 2nd + 3rd digits ; are twice the position of maximum camber in percent of chord

41. ♣ 4th + 5th digits ; section thickness in percent of the chord 2] example NACA 23012 ; ♣ 12(4th + 5th)% thick airfoil having a design of 0.3[2(1st digit)=0.3*10*2/3 or 2*1/10*3/2=0.3] and a maximum camber located 15% of c back from the leading edge. ♣ 2 ; =0.3*10*2/3 ; design lift coefficient=0.3 30 ; 15%*2=30%, twice position of max camber 12 ; max thickness in percent of the chord 3] Difference with 4 Digit Series - The NACA five-digit series uses the same thickness distribution as the four-digit series but the mean camber line is defined differently, however, in order to move the position of

42. maximum camber forward in an effort to increase . Indeed, for comparable thickness and cambers, the values for the five digit series are 0.1 to 0.2 higher than those for the four-digit airfoils. 3./ NACA 1-Series(Series 16) 1] Numbering System ; e.g. NACA 16-212 1st digit 1 ; series designation 2nd digit 6 ; location of minimum pressure in tenths of chord 3rd digit = 1st digit following the dash 2; design in tenth, so design lift coefficient = 0.2 4th + 5th digit ; the maximum thickness in percent of chord 2] Characteristics - The most commonly used 1-series airfoils have the

43. minimum pressure located at the 0.6c point and are referred to as series-16 airfoils. - The camber line for these airfoils is designed to produce a uniform chordwise pressure difference across it. - Operated at its design the series-16 airfoil produces its lift while avoiding low-pressure peaks corresponding to regions of high local velocities. - Thus, the airfoil has been applied extensively to both marine and aircraft propeller. 4./ NACA 6-Series 1] Numbering System ; e.g. NACA 651 - 212 a=0.6 1st 6 ; denotes the 6-series

44. 2nd 5 ; location of minimum pressure in tenth of chord for the basic thickness and distribution subscript 1 ; indicates that low drag is maintained at values of 0.1 1st digit after dash 2; design of 0.2 denoted by the 2 last two digits 12 ; specify the percentage thickness a=0.6 ; - The mean lines used with the 6-series airfoils have a uniform loading back to a distance of x/c=a. - Aft of this location the load decreases linearly. - The a=1 mean line corresponds to the uniform loading for the series-16 airfoils. - If the fraction, a, is not specified, it is understood to equal unity.

45. 2] Characteristics - The NACA 6-series airfoils were designed to achieve desired drag, compressibility and performance. These requirements are somewhat conflicting, and it appears that the motivations for these airfoils was primarily the achievement of low drag. - drag bucket

47. 6./ Low-Drag-Aerofoil Characteristics

49. General Aviation Airfoils ; GA(W) 1] definition and origin ; another family of airfoils evolved from the supercritical airfoils, but for low-speed applications 2] performance of GA(W)-1 airfoil ; see Fig.3.11,12,13 3] alternate notation for the GA(W)-1 = LS(1)-0417 where LS ; low speed 1; refers to a family 04 ; defines a design lift coefficient of 0.4 17 ; maximum thickness ratio in percent

50. 6./ Choice of Section 7./ Another Benefit of High Aspect Ratio 8./ Artificially Induced Laminar Flow 9./ Reducing Trailing-Vortex (Induced) Drag 10./ Improving Spanwise Lift Distribution Area Rule 11./ Wing-Tip Shape