Questions and Answers of Fundamentals Of Aerodynamics

What is the development trend of high lift devices in the future?
Air with a temperature of 25 °C will pass a thin plate longitudinally at a speed of 30 m/s. The pressure is atmospheric pressure. Calculate the thickness of the boundary layer at a distance of 200
The wings of a general aviation aircraft are rectangular (extended length 9.75 m, chord length 1.6 m). The aircraft flies at a constant speed at sea level at a cruising speed of 227 km/h. Assuming
In the case of the same Reynolds number Ret, try to find the ratio of the frictional resistance generated when water at 20 °C and air at 30 °C flow in parallel through a plate of length L.
Try to find the power to overcome air resistance when a car is driving at a speed of 60 km/h. It is known that the projected area of the car perpendicular to the direction of motion is 2 m2, the drag
Suppose a low-speed aircraft is flying at an altitude of 3000 m at an altitude of 360 km/h. If the wing area is 40 m2, the average chord length is 2.5 m. Try to estimate the frictional resistance of
In the wind tunnel, blow vertically to a disc with a diameter of 50 cm at a wind speed of 10 m/s, and try to find the resistance of the disc. The air temperature is 20 °C.
For a flat plate placed downstream in a two-dimensional incompressible flow, try the momentum integral method to find the wall friction stress and the friction resistance on one side of the plate F,
Please explain the physical meaning of the complete gas state equation.
What is the internal energy, pressure energy, kinetic energy, and enthalpy of a system?
Please write down the mathematical expression and physical meaning of the first law of thermodynamics in a static state.
Please explain the mathematical expression and physical meaning of the first law of thermodynamics in motion. Point out the difference with the steady flow energy equation of incompressible fluid
Write down the work done by the unit mass system to the outside world in the process of isovolume, pressure, and temperature.
Explain the physical meaning of adiabatic, reversible, and irreversible processes.
Write down the relationship between pressure and density in the adiabatic process.
What is the concept of entropy and why is dq/T a function of state?
For steady adiabatic flow (non-isentropic), please write down the relationship between the ratio of static pressure at any two points along the streamline and the Mach number.
For steady adiabatic flow (isentropic), please write down the relationship between the ratio of static pressure at any two points along the streamline and the Mach number.
What is the stagnation point parameter? Why does the total enthalpy remain unchanged and the total pressure decreases in the steady adiabatic friction pipeline flow?
Explain the physical phenomenon that the total temperature of a system remains unchanged and the total pressure decreases. Explain a physical phenomenon in which the total temperature and total
For an adiabatic steady flow in a sudden expansion pipe, please explain the relationship between the speeds at the sudden expansion section of unequal entropic flow, isentropic flow, and
Please derive the relationship between the relative increment of steady flow velocity and the relative increment of area in a variable cross-sectional pipeline?
Please derive the relationship between the ratio of the minimum cross-sectional area of steady flow to the cross-sectional area of any cross section and the Mach number in a variable cross-sectional
What is Mach wave? Explain the characteristics of Mach wave. Write down the pressure coefficient through Mach wave.
Write the continuous equation, tangential momentum equation, and normal momentum equation that pass through the Mach wave line?
Why is the Mach wave line emitted from the disturbance source when passing through the expansion angle? What wave is Mach wave?
What is the Mach wave when passing through the compression angle? Why does the Mach wave line converge?
Please explain the physical mechanism of shock wave formation. Why is the velocity of the shock wave greater than the velocity of the gas before the shock wave?
Why is the shock wave advancing speed less than the wave speed of the gas after the shock wave?
Write the governing equations (continuity equation, momentum equation, energy equation) through the shock wave.
Please write the expression of entropy increase through normal shock wave. Why is there an increase in entropy?
What is an oblique shock wave? Please point out the difference between oblique shock and normal shock.
Please explain how shock waves are generated in supersonic airflow.
Consider a room with a rectangular floor measuring 5 m by 7 m and a ceiling height of 3.3 m. The indoor air pressure and temperature are 101,325 Pa and 25 °C, respectively. Calculate the internal
Assuming that a Boeing 747 is flying at a standard altitude of 10,972 m, the pressure at a certain point on the wing is 19152 Pa. Assuming that there is an isentropic flow on the wing, calculate the
The pressure, temperature, and velocity at a certain point in the airflow are 101,325 Pa, 46.85 °C, and 1000 m/s. Calculate the total temperature and total pressure at that point.
An airplane flies at a standard altitude of 3048 m. The Pitot tube installed in the nose measures 105,336 Pa pressure. The plane is flying at subsonic speeds higher than 483 km per hour. The flow
An airplane flies at a standard altitude of 3048 m. The Pitot tube installed in the nose measures 106,294 Pa pressure. The plane is flying at subsonic speeds higher than 482 km per hour. Calculate
The temperature and pressure at the stagnation point of the high-speed missile are 245 °C and 790,335 Pa, respectively. Calculate the density at this location.
At the upstream of the shock wave, the air temperature and pressure are 14.85 °C and 101,325 Pa, respectively; just downstream of the shock wave, the air temperature and pressure are 417 °C and
Consider the water pour point where the flow velocity and temperature are 396 m/s and −6 °C, respectively. Calculate the total enthalpy at this time.
In the supersonic wind tunnel, the speed is negligible and the temperature is 727 °C. The nozzle outlet temperature is 327 °C. Assuming an adiabatic flow through the nozzle, calculate the exit
(a) Suppose that a 300-m-long pipe is filled with air at a temperature of 47 °C, and one end of the pipe generates sound waves. How long does it take for the wave to reach the other end?(b) If a
Suppose an airplane is flying at a speed of 250 m per second. If it is flying at a standard altitude of (a) sea level, (b) 5 km, and (c) 10 km, calculate its Mach number.
Assuming a point in the airflow, the pressure and density are 70927.5 Pa and 0.98 kg/m3, respectively. (A) Calculate the corresponding value of the isentropic compression coefficient. (B) Calculate
As shown in the figure, consider a wedge with a half angle of 15° in a water flow with a Mach number of 5. Calculate the drag coefficient of this wedge. (Assume that the pressure on the substrate is
Consider the separated curved bow shock wave in front of the two-dimensional parabolic bluff body shown in the figure. Free flow is Mach 8. Consider the two streamlines that pass through the impact
Consider the isentropic compression surface drawn in the figure. The Mach number and upstream pressure of the wave are M1 = 10 and p1 = 1atm (1atm = 101325 Pa), respectively. The water flows through
Please briefly describe the definition of airfoil thickness, camber and angle of attack and point to the thickness, camber and x value of the maximum camber of NACA4412 and NACA23012 airfoils.
The lift force affecting the airfoil is the incoming flow velocity V∞ and density ρ∞, chord length b, sound velocity a∞, thickness C, curvature f , angle of attack α, sideslip angle β, and
Explain the steady flow around a low-speed airfoil. What is the trailing-edge condition.
Explain the formation process of starting vortex and attached vortex.
Explain Kutta–Joukowski rule of lift circulation.
Try to point out the factors affecting the strength of attached vortex around steady airfoil and the reasons.
Try to establish the relationship between the strength of the attached vortex and the lift coefficient.
What is potential flow superposition for steady airfoil flow?
Describe the basic characteristics of surface source and surface vortex?
How to deal with the problems of thickness, camber, and angle of attack in the steady flow around a thin airfoil?
How to solve the lift problem of thin wing theory? Try to point out the variation trend of surface vortex intensity along the chord line.
Try to point the lift line slope, zero-lift angle of attack and zero-lift moment coefficient of flow around thin airfoil.
Try to point out the physical cause of leading-edge suction.
Briefly describe the solution idea of airfoil thickness problem.
What are the pressure center and pneumatic center and their relationship?
As the angle of attack increases, why does the moment around the aerodynamic center remain unchanged?
How does the lift coefficient of the airfoil change as the angle of attack increases? Why does the maximum lift coefficient of the airfoil appear?
Try to explain the physical parameters that the Re number affects the flow around the airfoil and the reasons.
For the steady flow around a flat plate airfoil without thickness at small angle of attack, try to prove that the concentrated vortex is arranged at 1/4 chord and the boundary condition is satisfied
What are the main parameters affecting the drag coefficient of airfoil steady flow?
Try to explain the physical causes of viscous pressure drag of steady airfoil flow.
Try to explain the main factors affecting the lift-to-drag ratio of an airfoil.
Try to point out the physical phenomenon of flow separation around airfoil. What are the trailing-edge separation and the leading-edge separation?
Briefly describe the effect of separation on airfoil lift and drag coefficient.
NACA 2412 airfoil with chord length of 0.64 m in the air flow under standard sea-level conditions is considered. The free flow velocity is 70 m/s, and the lift per unit chord is 1254 N/m. Try to
The vortices are distributed on the position of 1/4 chord of a plate at small angle of attack as shown in the fellow figure. Try to prove that if the boundary condition is satisfied at 3/4 chord
NACA 2412 airfoil with chord length of 0.64 m in the air flow under standard sea level conditions is considered. The free flow velocity is 70 m/s, and the lift per unit chord is 1254 N/m. Try to
Consider a thin flat plate with an angle of attack of 5°. Calculate: (a) the lift coefficient, (b) the moment coefficient about the leading edge, (c) the moment coefficient about the 1/4 chord
For the flow around a flat plate airfoil at a small angle of attack, try to prove that y(0) can have the following two forms of solutions.(1) (not met the trailing-edge condition)(2) (met the
NACA 23012 is considered in the study at α = 4°, while the CL = 0.and Cm,c/4 =−0.005. The zero-lift angle of attack is −1.1°. In addition, at α =−4°, Cm,c/4 =−0.0125. Based on the given
Thin symmetrical airfoils with an angle of attack of 1.5° are considered. According to the calculation results of thin airfoil theory, try to calculate the lift coefficient and moment coefficient of
People often ask such a question: can the wing fly backwards? To answer this question, perform the following calculation. Consider airfoils with a zero-lift angle of −3°. The lift slope is 0.1 per
The front section of an airfoil is a flat plate, and the rear section is a flat plate flap with a downward deviation of 15°, as shown in the fellow figure, try to calculate the CL at angle of attack
Consider a finite wing with an aspect ratio of 8 and a taper ratio of 0.8. The airfoil profile is thin and symmetrical. Calculate the lift and induced drag coefficient of the wing at an angle of
Given a trapezoidal swept-back wing, the wingtip chord length b1, the wing root chord length b0, the taper ratio η = b0/b1, and the spanwise length l, try to derive the calculation formulas of area,
Consider a rectangular wing with an aspect ratio of 6, an induced drag coefficient δ = 0.055, and a zero-lift angle of attack of −2°. When the angle of attack is 3.4°, the induced drag
A light single-engine general aircraft has a wing area of 15 square meters, a wingspan of 9.7 m, and a maximum gross weight of 1111 kg. The wing adopts NACA 65–415 airfoil, the lift coefficient is
Consider a finite wing with an aspect ratio of 6. Suppose the lift is distributed in an elliptical shape. The lift slope of the airfoil profile is 0.1/degree. Calculate and compare the lift
When the angle of attack is small, the polar curve of the wing can be expressed as a parabola try to prove CD = CDO+C
Explain the general principles of aerodynamic design for slender fuselage of transport aircraft.
Give the linearization assumption and governing equation of slender axis-symmetric body.
Explain the basic idea of cross-flow theory of slender body flow at high angle of attack.
Explain the physical concept of wing-body interference in the wing-body configuration flow.
Briefly describe the separation phenomenon of slender axis-symmetric body flow at high angle of attack.
What is subsonic flow? Why is the subsonic flow more curved perpendicular to the airfoil than the incompressible flow?
Derive a differential equations that describes the motion of an ideal compressible fluid.
Why the governing equation of ideal compressible potential flow is nonlinear?
What is the linearized equation for the velocity potential function of compressible potential flow? Please describe the linearization conditions.
Under the condition of linearization, please write out the formulation of the solution for the steady compressible potential flow? And A linearized expression of the wall pressure coefficient?
How to use affine transformation to solve the linearization of velocity potential function of steady compressible plane potential flow?
How to use affine transformation to transform the equation below to Laplace’s equation. Required Y = y, Z = z. 24 ax + a + ay a az = 0
Explain the difference between Karman-Qian formula and Prandtl-Glauert law Cp(Ma, a, c, f) = Cp(0, , c, f) Ma 1-Ma+ Cp(0.a.c.f) 1-Ma+1 2 IN

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