Questions and Answers of Fundamentals Of Aerodynamics

Why is the coefficient lift of a cambered aerofoil " 0 "?
What is induce drag?
Why is it said that the induced drag is part of lift?
State the effects of \(A R\) on coefficients of lift and drag.
Does the center of pressure change with airspeed and AoA?
Does fuselage produce restoring moment in rolling? Is it a stable feature to lateral stability?
Does airspeed affect pitch moment about the AC?
Does AoA affect the coefficient of pitch moment about the AC?
By using the diagram of yawing moment coefficients, explain the feature of sweep-back wing on directional stability.
By using the diagram of rolling moment coefficients, explain why an airplane is laterally stable if its rolling moment coefficient is positive when the sideslip angle \(\beta\) is negative
Find the aerodynamic center and the CP of RAF 15 aerofoil. The chord is \(1.8 \mathrm{~m}\). Its aerofoil data are in the table below, and its \(C_{M, A C}=-0.06\). Use the pitching moment about the
The Mach number of an aircraft is \(M=0.9\), and the local temperature is \(T=-10^{\circ} \mathrm{C}\). What is its airspeed?
Estimate TAS if an aircraft is at ALT \(=9500 \mathrm{~m}\) and its Mach number \(M\) is 0.5 .
An aircraft is in flight, and its \(\mathrm{TAS}=220 \mathrm{~m} / \mathrm{s}\). The ambient temperature \(T=245 \mathrm{~K}\). A local airspeed over an aerofoil reaches \(314 \mathrm{~m} /
Can you calculate the local speed of sound if you know ALT only? Is it true that the speed of sound definitely increases if both the air pressure and density are increasing?
Give flight situations in which both an aircraft's IAS (indicated airspeed) and Mach number will not change.
An aircraft is in flight, and its \(\mathrm{TAS}=220 \mathrm{~m} / \mathrm{s}\). The ambient temperature is \(T=253 \mathrm{~K}\). What is the stagnation temperature on its leading edge?
The Mach number of an airflow is \(M=0.7\). What is its speed coefficient \(M^{*}\) ?
Air flows through a converging-diverging nozzle. The Mach number of airflow at the inlet of the nozzle is \(M_{1}=0.75\), and the air condition is that at sea level at the inlet. The Mach number of
The Mach number before a normal shockwave is \(M=1.4\). What is the Mach number after the normal shockwave?
An airflow travels at \(M=1.5\) at sea level conditions before a normal shockwave. Find out the air property and airspeed after the shockwave.
The following table gives the values of the drag coefficient of a thin aerofoil at different free-stream Mach numbers. Calculate the drag of the aerofoil at sea level condition with these Mach
How do shockwaves move over an aerofoil as the free-stream Mach number increases from \(M_{\text {crit }}\) to \(M_{\text {det }}\) ?
Explain why \(\mathrm{CP}\) moves over an aerofoil when the free-stream Mach number increases from \(M_{\text {crit }}\) to \(M_{\text {det }}\).
What causes the significant rise in the drag coefficient as the freestream Mach number increases?
What is a bow shockwave? Explain the difference between the Mach number at the "Nose" of the bow shockwave and the free-stream Mach number.
What is Mach tuck? What is the adverse "stick force"?
What is the Mach buffet?
What is the difference between coffin corner and cross-over altitude?
Why does supercritical aerofoil produce less transonic drag?
When a supersonic airflow, \(M=1.8\), passes through a normal shockwave under sea level conditions, what are the values of the stagnation pressure before and after the normal shockwave?
The Mach number of an airflow is \(M_{1}=1.4\), air pressure \(p=1.012 \times 10^{5} \mathrm{~Pa}\), and air temperature is \(273 \mathrm{~K}\). This airflow, then, passes through an oblique
The Mach number of an airflow is \(M_{1}=1.4\), air pressure \(p=1.012 \times\) \(10^{5} \mathrm{~Pa}\), and air temperature is \(273 \mathrm{~K}\) before an expansion region. The deflection angle of
The airflow passes a twice-deflected surface. The free-stream Mach number, \(M_{1}\) is 2 at sea level, and the air flows along the path with a \(16^{\circ}\) deflection angle passes through a weak
Find the extreme value of \(f(x)=x^{2}+1\).
For function \(f(x, y, z)=x y z\), find \(\frac{\partial f}{\partial x}, \frac{\partial f}{\partial y}, \frac{\partial f}{\partial z}\) and \(d f\).
Integrate \(\int d p+\int v d v\).
Repeat Example 3 .15 using Wagner’s model of Example 3 .12. How similar are the predictions of Theodorsen’s and Wagner’s theory?Data from Example 3 .15Calculate the aerodynamic load responses
Repeat Example 3 .16 using Wagner theory. Try airspeeds higher than U∞ = 50 .47;how similar are the time and frequency responses predicted by Wagner theory and Theodorsen theory at such
Throughout Section 4.2, the pitching moment is calculated using the distance between the bound vortices and a datum (quarter chord or pitching axis). What differencewould it make if the pitching
Repeat Example 4 .8 with increasing amounts of camber. Is there a visible change in the frequency responses of the unsteady aerodynamic loads?Data from Example 4.8Calculate the aerodynamic loads
Develop frequency-domain versions of the SVPM and VPM techniques. Replace the wake model by the one used in Example 4 .8. Compare the frequency responses of the aerodynamic loads of a pitching and
Implement the non-circulatory lift calculation of Section 4.6 and use it to draw Figure 4 .40.Data from Section 4.6............................Data from Figure 4.40 Theodorsen's function was
Write a lifting line code to calculate the steady flow around a rectangular wing with any aspect ratio, taper twist and airfoil section.
Repeat Example 5 .10 for the other two wings tested by Queijo et al. (1956) and compare the VLM predictions to the experimental measurements.Data from Example 5.10Calculate the roll stability
Repeat Example 5 .4 using the SDPM and compare the resulting pressure distributions to the experimental pressures given in Kolbe and Boltz (1951).Data from Example 5.4Use the VLM to calculate the
Model the impulsive start of a wing with an elliptic planform using the SDPM and compare to the VLM and Jones/unsteady lifting line solutions.
Repeat Example 5 .7 using the SDPM. Are the differences in the predictions of the two methods more important for the lift or for the thrust/drag? Do these differences reduce as you reduce the
Repeat Example 5 .12 using the VLM. Does the VLM predict higher or lower aerodynamic loads than the SDPM?Data from Example 5.12Use the SDPM to simulate the plunging flexible wing experiments
Repeat Example 5 .13 for all Mach numbers tested by Lessing et al. (1960) using the incompressible SDPM code of the same example. How bad do the incompressible predictions get at Mach numbers of 0 .7
Implement the bending and twisting deformation of the LANN wing in Example 6 .3.Does the agreement between the compressible SDPM predictions and the experimental measurements improve when including
Repeat Example 6 .5 for the pitching oscillation test case at k = 0 .262 and M∞ = 0 .498 reported in Zwaan (1982). How good is the agreement between the SDPM and experimental oscillatory pressures?
Repeat Examples 6 .6 and 6 .7 for the M∞ = 1 .1 test case reported in Tijdeman et al. (1979b). Are the Mach box and Mach panel method predictions as good as for the M∞ = 1 .32 case? Are there any
Calculate the flow around the circular cylinder of Example 7 .4 using the vortex panel method of Example 7 .5.Data from Example 7.4Simulate viscous flow around a circular cylinder for Re = 105 using
There is a flat plate with a base area of 5 kg, sliding along the slant at an angle of 20° with the horizontal plane 60 cm times 40 cm × 60 c m times 40 cm. The thickness of the oil layer between
At a given point on the airfoil surface with very low speeds, the pressure coefficient is 0.3. If the Mach number of the free stream is 0.6, then CP calculated at this point is?
This is a two-dimensional parallel channel with a corrugated plate (sine curve), which the distance between the upper and lower walls is h, Prove that is a solution to the subsonic linearization
At low speed Incompressible flow, the pressure coefficient at a given point on the wing is 0.54. When the free flow Mach number is 0.58, by use of(1) Prandtl-Glauert law(2) Karman-Qian law(3) Laitone
A two dimensional airfoil is placed in the air so that its lowest pressure point appears on the lower surface. The pressure coefficient at this point is 0.782 when far field Mach number is 0.3. Try
The experiment of airfoil NACA006 was carried out in subsonic wind tunnel, and the slope of lift curve d at a = 0 was measured as Try to plot experimental curve and compare with the results from
Figure below shows four cases of flow over the same airfoil, where M∞ gradually increases from 0.3 to Mcr = 0.61 Point A on the airfoil is the minimum pressure point on the airfoil (and hence the
What is supersonic flow around airfoil?
What are the main characteristics of supersonic flow around airfoil compared with subsonic flow around airfoil?
Under the linearized condition, please write down the formulation of the definite solution problem of the velocity potential function of supersonic flow? The linearized expression of the wall
Please write down the general solution form and its physical meaning of the definite solution problem of the velocity potential function of supersonic flow?
For small perturbation supersonic flow around a thin airfoil, write the linearized decomposition of the surface pressure coefficient?
Please explain the main differences and pressure distribution of supersonic and subsonic flow around a plate?
Draw a diagram of supersonic flow around bending plate at 0 angle of attack and the pressure distribution on the upper and lower surface?
Write down the expressions of the lift coefficient and drag coefficient of supersonic flow around a thin airfoil, and state the physical meaning of each term?
Why does the lift coefficient of supersonic flow around airfoil decrease with the increase of inlet Mach number?
Write down the expressions of shock wave drag supersonic flow around a thin airfoil, and explain the physical meaning of each term?
Explain the relationship between minimum zero-lift shock wave drag and Mach number.
The Lockheed F-104 supersonic fighter, as shown in the figure, is the first fighter designed for sustained flight at Mach 2. The F-104 embodies good supersonic aircraft design. The airfoil thickness
Try to prove that the lift force in region I of the rectangular flat wing is equal to half of the lift force generated by the two-dimensional value of the supersonic velocity in this region. I Ma >1
What is transonic airfoil flow?
Compared with subsonic airfoil flow, what are the main characteristics of transonic airfoil flow?
What are the lower critical Mach number and the upper critical Mach number?
Write the expression of critical pressure. How to calculate the lower critical Mach number of transonic airfoil flow?
Try to sketch the trend of critical Mach number with the angle of attack, curvature, and thickness.
Try to sketch the trend of the lift coefficient of an airfoil with the increase of Mach number in the transonic region (Mach number 0.8–1.6)?
What is a shock stall? How does the lift coefficient change with Mach number when a shock stall occurs?
What conditions is the shock wave drag of transonic airfoil the largest?
In transonic airfoil flow, how to control the strength and stability of shock wave on the upper airfoil?
What are the main characteristics of the supercritical airfoil flow? Why can the flat upper wing suppress shock and reduce shock intensity?
In transonic airfoil flow, how can the lift loss due to the flatness of the upper airfoil be compensated by the change of the lower airfoil?
Why can variable camber suppress the shock wave on the upper wing of transonic flow?
Please point out the main measures to increase the drag divergence Mach number?
When Ma∞ of an airfoil increases to 0.8, the velocity of the maximum velocity point on the airfoil reaches the sound velocity. What is the pressure coefficient of the airfoil at the maximum speed
Try to calculate the lift line slope of a rectangular wing with aspect ratio λ = 10 in the condition of Ma∞ = 0.6 and compare it with the of the same wing in incompressible flow. a
The data obtained from the wind tunnel test of a rectangular thin wing of λ = 3 are as follows: On the basis of the above experimental data, try to calculate the lift line slope value of the
Try to calculate the lift coefficient of a thin rectangular wing of λ = 5 with Ma∞ = 0.85 using Prandtl–Glauert rule and the affine combination parameter chart method.
For the existing transonic flow = 0.95, a rectangular airfoil with relative thickness t = 0.08 and aspect ratio λ = 4. If the flow is similar try to calculate the aspect ratio and the relative
For the two-dimensional transonic flow, in the range of small disturbance theory, the flow can be considered as irrotational under the following conditions: Try to prove that if the above
Please explain the main characteristics of flow around multi-element airfoil and the physical mechanism of increasing lift?
The physical reason for the increase of lift coefficient of leading-edge slat? What is the effect of slot flow on lift increase?
Explain the effect of the downward deflection of the trailing-edge flap on the increase of lift?
What is the main difference between a drooped nose and slat flow?
The main physical causes of aerodynamic noise produced by multi-element wing?
Sketch the technical characteristics of the combined control of hinged flap with deflection of spoilers.
What is the development trend of high lift devices in the future?

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