Assume you have a supersonic airfoil at M_∞=1.5 If the airfoil has cl = -1.14723, what is the an...
- Assume you have a supersonic airfoil at
M_∞=1.5
- If the airfoil has cl = -1.14723, what is the angle of attack α in radians and degrees?
- What is the wave drag coefficient c_dw?
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Assume you have a supersonic airfoil at M_∞=1.5 If the airfoil has cl = -1.14723, what is the an...
a) There is a NACA 2412 airfoil with a chord length of 6ft. It has a...
a) There is a NACA 2412 airfoil with a chord length of 6ft. It has a 5 degree angle of attack during a sea-level flight at a velocity of 100 ft/s. Calculate the lift and moment (about the quarter chord) per unit span b) What is the profile drag coefficient for the airfoil in part a), as well as the drag force per unit span You are given the angle of attack. It is 5 degrees
2. A supersonic airfoil has a parabolic symmetric shape for upper and lower surfaces such that th...
2. A supersonic airfoil has a parabolic symmetric shape for upper and lower surfaces such that the maximum thickness is T at x =c/2, Compute the drag coefficient at zero incidence using thin airfoil theory. Compare with a diamond shaped airfoil with the same thickness Answers: M.-13e (ii) 33% greater than the diamond airfoil
2. A supersonic airfoil has a parabolic symmetric shape for upper and lower surfaces such that the maximum thickness is T at x =c/2, Compute the...
Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear supersonic theor...
Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data: • Total Wing Area: 18.22 m2 • Aircraft Mass: 8,000 kg -Is supersonic linear theory valid for this flight regime? Explain your answer.
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat pl...
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear a) supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data Total Wing Area: 18.22 m Aircraft Mass: 8,000 kg Is supersonic linear theory valid for this flight regime? Explain...
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angl...
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angles, θ¡et and θaf, at the leading and trailing edges, as defined in the figure and given in the table below. The airfoil operates at an angle of attack, α, and it is moving through air at a supersonic speed, M. The atmospheric temperature and pressure far upstream of the airfoil are T and p as specified in the table The various flow regions are...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow just outside the boundary layer if the freestream speed Vo = 120 m/s? • What is the local Mach number at this point if the altitude is 12 km? • What is the approximate value of C, for these conditions? Ans: V = 204.35 m/s, M = 0.693, CL ~ 1.15. -1.00 .. .... -0.50 .... ..... 0 0.1 0.2 0.3 0.4 0.5 0.6...
A thin straight wing with no twist has thin cambered airfoil with a0 -0.035 rad and...
A thin straight wing with no twist has thin cambered airfoil with a0 -0.035 rad and aspect ratio of 8. The wing is put at an angle of attack of 0.1 rad. Applying the fundamental lifting line equation for this wing at two different spanwise locations using two odd-term Fourier-sine series expansion for general circulation distribution I'(0) 2bV2£4, sin n0 yields the following set of equations: a 3.69.413.74 A-a0 a 6.73.4-3.624 -a0 a. From this analysis, determine the lift coefficient...
A very thin flat plate "airfoil" with a 1m chord is placed at an angle of attack α with respect t...
A very thin flat plate "airfoil" with a 1m chord is placed at an angle of attack α with respect to the free stream velocity Voo. The pressure distribution on the top and bottom surfaces of the "airfoil" are given by: pu -5.4x104 + 4x104(x - 1)2 (Nm2) pi- 1.7x105 + 2x104(x - 1)2 (Nm2) where X is the distance from the leadıng edge measured in meters. Neglecting shear stresses, determine the lift and drag forces per unit span. At...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
2. A supersonic airfoil has a parabolic symmetric shape for upper and lower surfaces such that the maximum thickness is T at x =c/2, Compute the drag coefficient at zero incidence using thin airfoil theory. Compare with a diamond shaped airfoil with the same thickness Answers: M.-13e (ii) 33% greater than the diamond airfoil
2. A supersonic airfoil has a parabolic symmetric shape for upper and lower surfaces such that the maximum thickness is T at x =c/2, Compute the...
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear a) supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data Total Wing Area: 18.22 m Aircraft Mass: 8,000 kg Is supersonic linear theory valid for this flight regime? Explain...
Question 2.10 A two-dimensional wedge-shaped airfoil, with chord, c, consists of straight-line segments with wedge angles, θ¡et and θaf, at the leading and trailing edges, as defined in the figure and given in the table below. The airfoil operates at an angle of attack, α, and it is moving through air at a supersonic speed, M. The atmospheric temperature and pressure far upstream of the airfoil are T and p as specified in the table The various flow regions are...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow just outside the boundary layer if the freestream speed Vo = 120 m/s? • What is the local Mach number at this point if the altitude is 12 km? • What is the approximate value of C, for these conditions? Ans: V = 204.35 m/s, M = 0.693, CL ~ 1.15. -1.00 .. .... -0.50 .... ..... 0 0.1 0.2 0.3 0.4 0.5 0.6...
A thin straight wing with no twist has thin cambered airfoil with a0 -0.035 rad and aspect ratio of 8. The wing is put at an angle of attack of 0.1 rad. Applying the fundamental lifting line equation for this wing at two different spanwise locations using two odd-term Fourier-sine series expansion for general circulation distribution I'(0) 2bV2£4, sin n0 yields the following set of equations: a 3.69.413.74 A-a0 a 6.73.4-3.624 -a0 a. From this analysis, determine the lift coefficient...
A very thin flat plate "airfoil" with a 1m chord is placed at an angle of attack α with respect to the free stream velocity Voo. The pressure distribution on the top and bottom surfaces of the "airfoil" are given by: pu -5.4x104 + 4x104(x - 1)2 (Nm2) pi- 1.7x105 + 2x104(x - 1)2 (Nm2) where X is the distance from the leadıng edge measured in meters. Neglecting shear stresses, determine the lift and drag forces per unit span. At...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
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