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Final MAE 3303
Attendance Quizzes 16-25
| Question | Answer |
|---|---|
| Under what condition can a flow be approximated as quasi one-dimensional flow? | When the rate of change of cross-sectional area and the curvature are small enough for one component of the velocity vector to remain dominant over the other two components. (Area cross-section of the flow stream changes but flow is one-dimensional) |
| For quasi-one-dimensional subsonic flow, what will happen happen when the velocity and pressure of the cross sectional area is reduced? What will happen to the velocity and pressure if the cross sectional area is increased? | When cross-sectional area decreases, velocity increases and pressure decreases, and vice-versa. |
| What geometrical feature must be present in order to accelerate an initially subsonic flow to supersonic speeds in a duct with slowly varying area? | In order to go from subsonic to supersonic speeds, flow must go through a convergent-divergent nozzle. |
| If a flow becomes supersonic in a duct with slowly-varying cross-sectional area, where will the sonic line be located? | The sonic line is located at the throat of the duct if flow becomes supersonic in a slowly-varying cross-sectional area. |
| What is the maximum exit Mach number we can get with a convergent nozzle? | M = 1 |
| Is the sonic throat area constant in a steady isentropic duct flow? Is a throat always the sonic throat? Is the sonic throat always located at the throat? | Sonic throat area is constant in a steady isentropic duct flow. The throat is not always located at the sonic throat. This sonic throat area is always located at the throat. |
| Once flow becomes choked in a nozzle, why is it not possible to increase the mass flowrate by decreasing the back pressure? | When flow becomes choked in a nozzle, the mass flowrate cannot be increased by decreasing the back pressure because the flow properties at the throat and throughout the entire subsonic section of the duct become unaffected ("frozen"). |
| What parameters does the maximum mass flowrate that a CD nozzle can deliver depend on? | m_dot = pt x At x Ut = p* x A* x a* |
| Can a normal shock wave stand somewhere in the convergent portion of a convergent-divergent nozzle? | Normal shock waves stand somewhere in the divergent portion of a CD nozzle, because a normal shock wave must be supersonic, and the flow upstream of the shock must be supersonic, which is not possible in a convergent nozzle. |
| What is the exit pressure of a CD nozzle? What is the back pressure of a CD nozzle? | Exit pressure of a CD nozzle is the pressure at the exit plane of a channel. Back pressure of a CD nozzle is the ambient pressure created by a vacuum pump/valve or local atmosphere. |
| What will happen if the back pressure of a CD nozzle is increased to the value that is higher than the exit pressure predicted for a subsonic-supersonic isentropic flow? | When back pressure of a CD nozzle is increased to the value higher than the exit pressure, the CD nozzle exit flow is supersonic. |
| What is the perfect expansion nozzle back pressure condition? | Pb = Pe = Pe6 = Pis |
| What is an under expanded nozzle exit condition? What flow structure will develop downstream of the nozzle? | Pb < Pe6 = Pis Expansion waves develop downstream of the nozzle when nozzle condition is under expanded. |
| What is an over expanded nozzle exit condition? What flow structure will develop downstream of the nozzle? | Pe5 > Pb > Pe6 = Pis Shock waves will develop downstream of the nozzle when under over expanded conditions. |
| What is a diffuser? Does the flow velocity increase or decrease along a diffuser? Does the flow pressure increase or decrease along a diffuser? | Mechanisms to slow high speed flows to low speed flows are called diffusers. Velocity decreases along a diffuser, while pressure increases. |
| If there are multiple sonic throats in a steady and adiabatic channel flow, what is the relation between stagnation pressures and sonic throat areas? | Their relation is constant. |
| What is the definition of irrotational flow? | Flow where the value of vorticity is zero. |
| What kind of flows would the velocity potential equation be exact? | The velocity potential equation is for steady, inviscid, potential and isentropic flow. |
| For incompressible flow, is the velocity potential equation a linear or non-linear partial differential equation? For compressible flow, is the velocity potential equation a linear or non-linear partial differential equation? | The velocity potential equation is linear for incompressible flow, and non-linear for compressible flow. |
| What type of differential equation is the linearized velocity potential equation for subsonic flow? What type of differential equation is the linearized velocity potential equation for supersonic flow? | The partial differential equation is elliptical for the velocity potential equation of subsonic flow, and hyperbolic for the velocity potential equation for supersonic flow. |
| At what conditions will the linearized velocity potential equation be reasonably valid? | The linearized velocity potential equation is reasonably valid for subsonic or supersonic flow over a thing body at a small angle of attack. |
| State the formula for linearized pressure coefficient. | Cp = -2(û/V_freestream) |
| What is the Prandtl-Glauert compressibility correction for pressure coefficient? | Look at attendance quiz, I can't really type this out. |
| What is the utility of the Prandtl-Glauert similarity rule? | The Prandtl-Glauert similarity rule provides a simple correction factor to relate pressure coefficients in incompressible flow to those in subsonic compressible flow, enabling easier aerodynamic analysis at moderate Mach numbers. |
| Why is it not possible to use a simple transformation to connect linearized supersonic flow to incompressible flow? | Transformation to connect linearized supersonic flow to incompressible flow is not possible because supersonic flow is hyperbolic and subsonic flow is elliptical. |
| According to the linearized supersonic flow theory, both oblique shock and expansion waves are treated as what kind of wave? | According to the linearized supersonic flow theory, both oblique shock waves and expansion waves are treated as weak, linear perturbation waves. |
| The pressure coefficient in linearized supersonic flow is proportional to what geometrical feature of a body? | The pressure coefficient is proportional to the upper and lower surfaces of a body. |
| According to the supersonic linearized theory, the lift coefficient of an airfoil is related to what geometrical features? | Angle of attack for a given mach number. |
| According to the supersonic linearized theory, the wave drag coefficient of an airfoil is related to what geometrical features? | The wave drag coefficient of an airfoil is related to all geometrical features (angle of attack, mean camber (z) and thickness (t)). |
| According to the supersonic linearized theory, the pitching moment coefficient of an airfoil is related to what geometrical features? | Angle of attack and mean camber line. |
| For the same thickness envelop and angle of attack, which airfoil has less wave drag, symmetric or cambered? | Symmetric |
| Where is the center of pressure located on a supersonic thin airfoil? Where is the aerodynamic center located on a supersonic thin airfoil? | The center of pressure and aerodynamic center are both located at the mid-chord. |
| Does the lift coefficient increase or decrease with increasing Mach number in subsonic flow? Does the lift coefficient increase or decrease with increasing Mach number in supersonic flow? | Lift coefficient decreases with increasing Mach number in subsonic flow, while increasing with increasing Mach number in supersonic flow. |
| What is the definition of the critical Mach number? Why is the critical Mach number of interest? | The critical Mach number is the free stream Mach number at which sonic flow is first achieved on the airfoil surface. It tells us the flight Mach number where locally supersonic flow exists. |
| Will the critical Mach number increase or decrease as the airfoil thickness increases? | The critical Mach number decreases as thickness increases. |
| What is the critical pressure coefficient of an airfoil? | The pressure coefficient at the point where the sonic flow is first achieved on the airfoil surface at the critical Mach number. |
| What is meant by drag-divergence Mach number? Why does the drag coefficient increase rapidly once the free-stream Mach number reaches the drag-divergence Mach number? | The drag-divergence Mach number is the Mach number at which the drag suddenly starts to increase rapidly. This rapid drag rise occurs because the formation of shock waves causes flow separation, and increased wave drag. |
| Typically by what factor does the drag increase as an aircraft passes through Mach 1? What is meant by "sound barrier"? Does it really exist? | The drag increase roughly by a factor of 10 as an aircraft passes through Mach 1. "Sound barrier" is the myth that sonic speed is the fundemental limit of aircraft flight speed. While existing, it does not pose to be a physical limit of flight speed. |
| For quasi-one-dimensional supersonic flow, what will happen happen when the velocity and pressure of the cross sectional area is reduced? What will happen to the velocity and pressure if the cross sectional area is increased? | When cross-sectional area increases, velocity increases and pressure decreases, and vice-versa. |
| What is meant by the drag divergence Mach number? Why does the drag coefficient increase rapidly once the free-stream Mach number reaches the drag-divergence Mach number? | The drag-divergence Mach number is the value of free-stream Mach number at which the drag begins to increase rapidly. The drag coefficient increases rapidly because there is an extensive region of supersonic flow is formed. |
| Typically by what factor does the drag increase as an aircraft passes through Mach 1? | Drag increases roughly by a factor of 10 as an aircraft passes through Mach 1. |
| What is meant by sound barrier? Does it really exist? | A sound barrier is the the myth that the fundamental limit of aircraft speed is Mach 1. While the pressure coefficient peaks at Mach 1, planes can move past this speed. |
| How is a supercritical airfoil distinguished from a convectional airfoil in geometry? What is the advantage of supercritical airfoil theory? | The supercritical airfoil has a relatively flat top, the forward 60 per-cent of the airfoil has negative camber, which lowers the lift. To compensate, the lift is increased by having extreme positive camber on the rearward 30 percent of the airfoil. |
| How will the quarter-chord pitching moment coefficient for a supercritical airfoil compare to a conventional airfoil? | The quarter-chord pitching moment of the supercritical airfoil is significantly higher than for a conventional airfoil. |
| State the area rule for transonic flow. Does the rule require widening or thinning of the fuselage in the vicinity of the wing? | To reduce the peak drag near Mach 1, the variation for cross-section area for an airplane should be smooth, with no discontinuities. This requires the thinning of the fuselage. |
| By what factor can the peak drag near Mach 1 be reduced through the use of area rule? | Peak drag near Mach 1 be reduced through the use of area rule by a factor of 2. |
| In general are Mach lines straight lines or curved lines in supersonic flow? | Mach lines are straight lines. |
| Characteristic lines are Mach lines? True or false? | True |
| How many characteristic Mach lines are at a point in supersonic flow? | Two characteristic Mach lines, I and II. |
| What is the angle between the characteristic line and the flow direction at a point in supersonic flow? | μ |
| What quantities are constant along characteristic lines? | v and θ. |
| What are the two approaches to calculate steady 2D isentropic irrotational supersonic flow properties using the method of of characteristics? | 1. Point to Point 2. Region to Region |
Created by:
philippsand
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