What exactly is an ATC clearance and how is it requested? What exactly does it mean to get an ATC clearance?
I’m asking the question because I read point 8015 of the Standardised
European Rules of the Air (SERAs), which states that ATC clearances are requested by submitting a flight plan. I always thought that this was done by a radio-request when asking for start-up (at least for IFR-flights)...
And what does it exactly mean to get certain ATC instructions during the flight (for example in the case of vectoring)? Are these amendments to the initial clearance or do I mix something up?

I realize that the question was asking about EASA, but perhaps an FAA perspective will increase understanding.  The FAA Pilot Controller glossary defines clearance as:

AIR TRAFFIC CONTROL CLEARANCE [ICAO]− Authorization for an aircraft to proceed under conditions specified by an air traffic control unit.

Note 1: For convenience, the term air traffic control clearance is frequently abbreviated to clearance when used in appropriate contexts.

Note 2: The abbreviated term clearance may be prefixed by the words taxi, takeoff, departure, en route, approach or landing to indicate the particular portion of flight to which the air traffic control clearance relates.

In short, clearance is permission, and is generally associated with operating under IFR in controlled airspace.
I will disregard taxiing at a tower controlled field because it can be done under IFR or VFR.  Additionally, I have heard compelling arguments that you are not “cleared” to taxi, but are given taxi “instructions”.  (Ground controllers will never use the word “cleared” when telling you to taxi).
But that is a separate discussion...
The process begins with filing an IFR flight plan.  This notifies ATC of your intentions, and gives them time to enter your request into the system.
The actual clearance is issued when you call them, (radio at towered field, land line at uncontrolled field) field and put your clearance “on request”.  When you call, ATC will issue the clearance.  It will be either what you filed for, or it may contain amendments.
In the clearance you may be told to expect radar vectors.  This then becomes part of your clearance.  However, vectors assigned as needed during flight for traffic separation aren’t generally considered an amendment to the clearance.  Think of them more as a temporary detour around something.
For example, if ATC needs to amend your clearance they will tell you specifically that they have an amendment, and ask when you are ready to copy.  The amendment will generally be a rerouting that you are expected to follow on your own, and may also include a different altitude.
If given a vector, however, the expectation is that is it a temporary heading needed by ATC to provide traffic separation.  Once the need for positive ATC control is past they will clear you to resume your own navigation along your flight planned route.


Can one use a calculator for the EASA part 66 exam? For the first 3 modules you will need to do some calculations, is a calculator allowed?

Calculators are not allowed during examination. Therefore all calculations should be feasible without a calculator. Where a question involves calculations not feasible without a calculator, such as root 10, then the question should specify the approximate value of root 10

I found the answer on page 7 of this pdf.


Do turboprop and piston engines perform differently at higher altitudes? When flying at higher altitudes, does it matter if you have a turboprop vs a piston engine driving a propeller?

In general, a turbine engine will produce more power than a piston engine at higher altitudes. A turbocharged piston engine, with compressed air intake, will lose less power at altitude than an normally aspirated piston engine. A turbine engine will lose even less.

As a rough example, a normally aspirated Cirrus SR-22 has a service ceiling of 17,500 ft. A turbocharged SR-22 has a ceiling of 25,000.

A turbocharged piston Piper Malibu Mirage has a service ceiling of 25,000 ft., while the turbine powered Malibu Meridian has a ceiling of 30,000 ft.

The turboprop Pilatus PC12 and Socata TBM-850 have service ceilings of 30,000 ft. or above.


Why must Vr be set so precisely? Why must VR (Rotation Speed) be set so precisely? I mean why can't they just say:

  Rotate the Aircraft at 120 to 150 knots

And then, the pilot thinks:

  OK, I have a full plane and half full tanks, I think I will rotate at around 135 knots

And then rotates at around that speed?

This answer is written for air transport category aircraft.
Introduction
During take-off there are three operationally significant speeds that ensure a safe take-off:

V1 - the take-off decision speed
VR - the rotation speed
V2 - the take-off safety speed

In addition there are three technically important speeds:

VMU the minimum unstick speed
VMCG the minimum controllable speed on the ground in case of an engine failure
VMCA the minimum controllable speed airborne in case of an engine failure

V1 is the take-off decision speed. Below V1, take-off can aborted safely without overrunning the end of the runway. Above V1 the aircraft is unlikely to be able to stop before the end of the runway so in case of an engine failure it is safer to get airborne. V1 needs to exceed VMCG to ensure the aircraft doesn't veer off the runway when accelerating to take off with a single engine.
VR is the rotation speed. At this speed the aircraft should be rotated and will get airborne soon after. The rotation speed ensures that, in the case of an engine failure, lift-off is possible and
V2 is reached at 35 feet at the latest. At the end of the rotation manoeuvre, the aircraft should be airborne. Otherwise the tail will hit the ground. Therefore, VR will need to exceed VMU.
V2 is the minimum speed that needs to be maintained up to acceleration altitude, in the event of an engine failure after V1. Flight at V2 ensures that the minimum required climb gradient is achieved, and that the aircraft is controllable. V2 speed is always greater than VMCA to ensure the aircraft is controllable in flight if an engine failure should happen.

Answer
What will happen if we just pick a VR that "feels good"?
If VRfeel is lower than VMU, we will end up with a tail strike. Even if the rotation would stop just before the tail strike, the increase in aerodynamic drag (lift is generated) will cause the take-off run to be longer than needed, risking not being able to climb over objects at the end of the runway. In addition, if the aircraft has an engine failure during rotation, it may not be able to reach V2 and could become uncontrollable and crash.
When VRfeel is too high, the aircraft risks exceeding the maximum safe tire speed. There is also a risk of not reaching VRfeel before the end of the runway.

References
Airbus operations briefing note on takeoff and departure operations; understanding takeoff speeds
Boeing article: exceeding tire speed
rating during takeoff


Wing vs wall to wall wing,total drag comparation at same AoA? Compare total drag of wall to wall wing vs wing(open tips) in wind tunnel,AT SAME AoA.
My first intuition is that wall to wall wing has higher pressure drag at outward parts of wing,because side-walls "act" as big endplates,so that makes difference.Reslut is higher total drag.
But

•wall to wall wing
If aerodynamic force1 =70N and angle A = 75° Drag1 =cos75 x 70N=18.1N
•wing
Aerodynamic force2 has lower magnitude lets say 60N and reduced effective AoA is 1° ,then angle B=74°  Drag2= cos 74 x 60N =16.5N
(even if I choose 65N(only 7% reduction in aerodynamic force)drag will be smaller at wing ,cos74 x65N=17.9N)
I think Peter Kampf has best opinion,this can not be prove with data because we dont know how much is magnitude of aerodynamic force at wing is reduced compare to wall to wall wing..
Wing has tilted  aerodynamic force back but in same time has lower magnitude ,so it cant be prove is drag higher or lower,it can go in both directions.
Vector tilt back at wing is usualy very small,maybe 1°,so my general oppinion is that vector tilt back has less contribution than reduction of aerodynamic force magnitude at wing...
What is your opinion?

I agree with your conclusion that the wall-to-wall wing has more drag than a free-flying wing at the same angle of attack.
But what is the point of this comparison?
People build wings to create lift. These wings might be on airplanes, but also on helicopters, propellers, the fin keels of boats or the blades in turbo machinery. Their purpose is in all cases to create lift. Not drag. Drag is an accepted price for that lift creation, but not their purpose.
So it is pointless to ask for a drag comparison when disregarding lift at the same time. This is akin to asking if a wing that is not moving creates no drag.
You don't even let us know the wingspan of those hypothetical wings. Are they in the same windtunnel, so the free-flight one has less aspect ratio, span and surface? Or is it the same wing, with the windtunnel removed?
So please stop repeating the same pointless and insufficiently specified question over and over. You will not get useable insights because the answers can not be helpful.


How to do blade element momentum theory analysis for different geometries of airfoil, e.g. NACA airfoils? I want to analyse NACA airfoils using BEMT. I have only yet done it for basic symmetric airfoils. I would like to know what all assumptions remain same in NACA airfoil analysis and to what degree it is accurate. Like for starters, will accounting for Cl-alpha change and geometrical changes in angle of attack be enough to analyse NACA airfoils?

The only difference would be the lift and drag coefficients used.  You'd need those to be specific to the airfoil(s) you choose.  Momentum theory is otherwise not sensitive to the airfoil shape.
Given the same lift and drag, momentum theory will give you the same induced velocity.


When is a flight plan required? Do all flights need to have filed flight plans or only ones that fit certain criteria? What are the criteria?

In the US you need to file a flight plan for the following flights:

IFR flights
Defense VFR Flights, which are required to fly in the ADIZ that is off the coast of the US (an IFR flight plan may also be used).
Some TFR's only allow air traffic to fly within their boundaries if they are on an IFR or VFR flight plan.
A DC FRZ flight plan is required to fly VFR in the DC FRZ.
A DC SFRA flight plan is required to fly VFR in the DC SFRA.
When entering or leaving the U.S., an ICAO international flight plan (IFR or VFR).

Other than that, no flight plans are required, but are strongly recommended so that search and rescue gets called if you don't show up at your destination on time!


Why do some airplanes allow separate right and left rudder pedal adjustment? Ideally, when a pilot adjusts the rudder pedals both pedals should move together, to have equal leg length for the left and right legs.

Some aircraft like the Dornier 228 have rudder pedal adjustments where you have to adjust the left and right rudder pedals separately. What is the logic of having provision for separate adjustment of the left and right leg length to reach rudder pedals?

As with any question along these lines—"Why did the designers do X"—a definitive answer is very difficult absent specific documentation, which is rare.

I am not familiar with the Dornier 228, but I am familiar with the King Air B200, which is a very similar class of aircraft. Beechcraft also designed the King Air with individual pedal adjustment.

In my examination of that system, I believe the primary reason is simplicity: the individual adjustment system on each pedal is simple and lightweight, whereas a system to adjust both pedals simultaneously would likely be heavier and more complex. Weight and complexity are both undesirable characteristics in aircraft design.

See this excerpt from the King Air B200 Illustrated Parts Catalogue (IPC) depicting the left hand side rudder pedal assemblies:

Source: KA B200 IPC 27-20-00-06

I have highlighted the rudder pedal adjustment lever assemblies—B and C—in light blue. I have highlighted the adjustment pin—No. 375—in green. (Note that the IPC incorrectly depicts the two adjustment levers in reversed locations; see the aircraft photo below.)

The figure clearly shows the simplicity of the design and the way the adjustment relates to the overall rudder pedal control system. The adjustment lever is a simple assembly of stamped metal and hardware, light and cheap to manufacture. The system is simple and intuitive to use and requires little or no rigging.

This is a photo of the left hand side rudder pedal assembly in an actual King Air B200; adjustment is accomplished by depressing the lever with the outer edge of one's shoe and moving the pedal forward or aft:

Any design allowing simultaneous dual adjustment through a single adjustment action would likely have resulted in some combination of increased complexity, expense, weight, or maintenance requirements.

Moreover the system, as designed, allows for separate adjustment, which could be helpful for some individuals or situations.


Term to differentiate between a flight and a flight instance? What's a term to differentiate between a flight schedule (QF 72) and a particular instance of that flight schedule (QF 72 departing on July 25)?

For example, I might ask "Which flight are you taking", and you might reply "QF 72". What I really meant was "Which scheduled instance of QF 72 are you taking?"

The flight schedule and a specific instance of that flight schedule are both referred to as a "flight". What are the terms that more precisely differentiate between these two concepts?

You might say "the QF 72 service between Singapore and Perth" for the flight schedule, and "the QF 72 flight on July 25" for a particular instance of the flight.

The term service used in this way is seen more often in other transportation sectors such as bus and train. But it can also apply to scheduled air services.


Why does the Zapata Flyboard Air UL consume so much more fuel/hr than the Mosquito Air? They are both ultralight and abide by FAA ultralight reg 14 CFR part 103.  The Mosquito Air is an ultralight 1 person helicopter whereas the Zapata Air UL is a micro jet turbine powered 1 person ultralight hoverboard. They weigh the same and both only carry 5 gallons of fuel as this is dictated by part 103.

The 5 gallons will last 1 hour of cruise flight time on the Mosquito while only 8 minutes on the Zapata.  Granted we are dealing with different technologies here and different fuels (Zapata uses kerosene, Mosquito uses avgas) I was surprised by the fuel burn difference.

The bare physics tells us that equal masses should require the same amount of thrust, thus energy, to hover in place where lift equals weight.  Kerosene has similar energy density to avgas.  And, a jet turbine is supposed to be much more efficient than a 2-stroke, 2 cylinder gasoline engine.

The Mosquito boasts of a 64hp engine "with the highest power to weight ratio on the market today".  The Zapata company seems not so much concerned with the gas guzzling fuel consumption.  The 2 companies seem to have different priorities here, so I'm guessing the fuel efficiency of the Zapata could be improved.  Still, this is a huge difference that is difficult to understand.

It's the difference between jet engines and propellers. It is a lot more efficient to accelerate more air to a low velocity, than it is to accelerate less air to a higher velocity. For fixed wings, the only reason to use jets is if propellers hit their speed limit, when the tips break the speed of sound.

Hovering flight is particularly demanding on fuel consumption. The Mosquito Air burns through its fuel a lot faster than 1 hour when it stays in hover all the time. Cruising in a helicopter makes the fuel consumption  a bit more similar to fixed wing: the rotor blades start to act more like wings and can use airspeed to reduce induced drag. Still less efficient than fixed wing cruise, but more efficient than helicopter hover.

The Zapata Air UL is always in a hover, and has no wings, and is a jet. It never gets the increased efficiency effect of forward flight that a rotor gets. And jets are really very inefficient at low speeds!


What is the largest aircraft that has been flown inverted successfully? What is the largest aircraft that has been flown inverted? For the purpose of the question:

Sustained inverted flight is not necessary, therefore barrel rolls and loops count
The aircraft must be undamaged after the maneuver
The maneuver must be intentional (i.e. China Airliners 006 does not count)
Simulator runs do not count. The question is not "can it possibly", but "has anyone tried successfully".

The largest I am aware of is the Boeing 707, but that was more than 50 years ago. Has anyone attempted a similar feat in a larger aircraft?

Theoretically ANY airplane can be flown inverted, if you know what you are doing with it and make sure it is loaded properly throughout the maneuver.  Boeing test pilot Tex Johnson famously rolled the 707 prototype - twice - during a demo flight for airline executives at the SeaFair hydroplane races on Lake Washington in Seattle.  He nearly got fired for the stunt but the plane was perfectly fine.

The real issue is load factor: the total load imposed upon the wings divided by the gross weight of the aircraft.  Don't exceed theses values and have the aircraft flown within the required speed range and it doesn't care whether it's erect or inverted during the process.

In general though, unless an aircraft is specifically designed for inverted flight with inverted i.e. negative G fuel and oil systems, inverted flight must be maintained with a positive load factor to prevent fuel starvation and damage to the engines.

As a side note to this, DO NOT TRY AEROBATICS IF YOU HAVE NO FORMAL AEROBATIC TRAINING AND/OR ARE ATTEMPTING TO DO SO IN AN AIRPLANE IN WHICH THEY ARE NOT APPROVED.  It's true guys like Sean Tucker have done aerobatics in a Cessna Corvalis but those are experienced airshow demonstration pilots with over 20,000 hours of logged flight time and more experience at the edge of the envelope than you can shake a stick at.  You do not have this prerequisite experience and most likely will kill yourself attempting these maneuvers.


Why do some commercial aircraft have no landing gear cover? The B737 has no cover on its landing gear, as may be seen in this picture. Doesn't this increase the air drag?

Almost all other airliners have doors that close off the rubber of the wheels, for example this B787.

Why doesn't the B737 have a cover on its landing gear?

Yes, it does increase drag.  Anything that is not smooth, where the air flow will be disturbed, will generate drag.

The 737 was originally designed only for short haul and space is at a premium.  Adding doors and all the associated mechanisms adds complexity, weight, additional maintenance requirements and more possibilities for failure.

Boeing have had many opportunities to add doors and have chosen not to; so they, and their customers, do not see the drag penalty as more expensive than the extra weight and maintenance.


Does a plane have brakes to stop or slow down while flying? Does a plane have brakes to stop or slow down while flying?

For example, if the pilots see a large flock of birds at some distance in front of the airplane, and they want to reduce speed suddenly. What will happen if an airplane doesn't have such brake?

An airplane can slow down and reduce its speed while in flight. The easiest way to do so is to reduce the amount of thrust that the engines are producing. This will produce an almost immediate reduction of the airspeed, especially if the plane is maintaining the same altitude.

There are also devices called air brakes and spoilers that can be further used to reduce speed. These, however, are never1 used in normal, level flight by passenger2 aircraft and are normally only used to reduce speed during the descent and landing phases of the aircraft3.

If an airplane reduces its speed too much, it will of course stall and start dropping precipitously, at which time the airspeed usually also increases again. The slowest speed an aircraft can maintain at a given altitude without stalling is listed in its flight envelope.

However in the scenario you mentioned (obstacles ahead), it's usually easiest to simply fly around the obstacle.

1: Never say never on the internet.

2: The situation is different in military aircrafts, you can see the approaching Su-27 using its airbrake in this video

3: The wheels of most aircraft also have brakes, similar to those found on automobiles. Obviously these can only be used when the plane is touching the ground. Lowering retracted landing gear will also increase drag. There are also thrust reversers on some aircraft that are used in the final stages of landing. Use of thrust reversers during flight can be disastrous although some older aircraft (including the C-17A) were designed to use thrust reversal to scrub speed in-flight.

Finally, some gliders and larger aircraft have drag or drogue parachutes to help slow them down, for example for use on short runways.


What makes V1 one or two knots less than Vr on this performance table? (B738 AFM, 0-2000 feet elevation, up to 27°C) Example of V1 and Vr differences.

The table shows that the V1 speeds are invariably one or two knots lower than their matching Vr speeds. Given that V1 can equal Vr, what's the limiting factor (probably factors) that puts V1 one or two knots below Vr? Vmbe might be one, but is it the only possible limiting factor here?

Or perhaps the V1 speeds are the same with the matching Vmcg speeds because that's operationally more advantageous in some ways than V1s being the same as the higher Vr speeds?

Or are the V1 speeds the speeds that bring the balanced field lengths, which require the mimimum runway distance?

$V_R$ is:

FAR 23 (e.g. multi-engine small planes):

  $\geq 1.05\;V_{MC}$ (105% of minimum control in flight)

  $\geq 1.1\;V_{S_1}$ (110% of V-stall for a specific configuration)

FAR 25 (e.g. transport jets):

  $\geq V_1$

  $\geq 1.05\;V_{MC}$

  Must result in $V_2$ before 35 ft height

  Must result in $V_{LOF}$ [when the airplane is rotated at its maximum practicable rate]

    $\geq 1.1\;V_{MU}$ (all engines operative; AEO)

    $\geq 1.05\;V_{MU}$ (one engine out)

    $\geq 1.08\;V_{MU}$ (AEO, geometry limited)

Geometry limited is when there is a greater risk of a tail-strike, e.g., a stretched variant of an airplane (think A321 or 737-900).

So the 1-2 knot disparity in your example can be down to the difference between $V_{MCG}$ and the 105% of $V_{MC}$ or the different $V_{MU}$ speeds. (See below for full text on points 2 and 4.)

To clear any confusion about the Weight:V1 tables such as the one in the question, those tables are not for 1 runway length, those weights are the field limit weight, i.e., each higher weight in the table is a longer runway.

(Twitter) $V_{MU}$ test.

Source: http://wpage.unina.it/fabrnico/DIDATTICA/SPER_VOLO/APPUNTI_3CFU/11.2-TO&LAND.pdf

FAA AC No: 25-7C text:

  VR may not be less than V1; however, it can be equal to V1 in some cases.
  VR may not be less than 105 percent of the air minimum control speed (VMCA).
  VR must be a speed that will allow the airplane to reach V2 at or before reaching a height of 35 ft. above the takeoff surface.
  VR must be a speed that will result in liftoff at a speed not less than 110 percent of VMU (unless geometry limited) for the all-engines-operating condition and not less than 105 percent of the VMU (unless geometry limited) determined at the thrust/weight ratio corresponding to the one-engine-inoperative condition for each set of conditions such as weight, altitude, temperature, and configuration when the airplane is rotated at its maximum practicable rate.

(Emphasis mine.)


Is this one of the engines from the 9/11 aircraft? I ran across an interesting set of post-9/11 photos and I noticed this picture in particular:

The closest item on the truck looks a bit like an airplane engine, but I can't tell for sure. Is it one of the engines from the 9/11 aircraft?

They are definitely not jet engines. They look more like electric motors or generators or blower units or something.


What is the purpose of the extending wingtip panels on the Su-25? Is it to increase the air resistance of the aircraft?

Source: reddit.com

Those are airbrakes, they allow the aircraft to increase its drag substantially to bleed off excess energy. They are hydraulically operated, and need pressure for both deployment and retraction, i.e. they do not auto-stow, at least on this aircraft.

Airbrakes are a handy feature, especially for a ground attack aircraft like the Su-25 in your picture, as they help both with landing on short airstrips and with dive bombing, a tactic that requires to dive almost vertically on a target to minimize slant errors in targeting. This is made possible by the airbrake, which prevents the aircraft from exceeding its dive speed.

Other aircraft types use or have used airbrakes in the past, mainly to dissipate energy during an approach. Examples include some airliners, most fighters and nearly all gliders.

Fighters are usually optimized for high speeds which are dangerous to land at, and an airbrake enables them to quickly slow down while on final approach, remaining at a comfortable speed until the last moment. They also have some occasional uses in WVR (Within visual Range) combat, but this will be rabidly argued against by the BVR (Beyond Visual Range) and stealth cults.

Gliders also use airbrakes to control their speed on approach and when diving, but also to arrest undesired climbs, when caught in a thermal when they need to descend. Since typically gliders lack a hydraulic or even electric system, their airbrakes are actuated by the pilot's muscles and are more akin to spoilers.

The geometry of an airbrake can vary significantly, but the principle is the same: a large surface extended into the airstream that generates mostly drag. their location on the airframe is a compromise between structural, mechanical and control considerations, i.e., it needs to be strong enough to resist the loads, easy enough to route hydraulic lines or control rods to, and not significantly impact the aircraft's stability.  The examples below are a BAe 146, an F-15, a Slingsby Capstan and a MiG-23:


How can I calculate maximum rate of climb? By what way and with which variables could you determine a plane's maximum rate of climb per time? If I'm not mistaken, I'm looking for VY.

To calculate your possible climb speed $v_z$, you will need

Your engine's thrust $T$
Your airplane's drag $D$
Your airplane's mass $m$

Calculate how much power is needed to overcome drag, and any excess can be used for climbing: $$v_z = v\cdot sin\gamma = v\cdot\frac{T-D}{m\cdot g}$$

Note that this equation makes use of several simplifications, but works well for propeller and slow turbofan aircraft with moderate flight path angles $\gamma$.

To do this with more precision, you need to account for the fact that the aircraft should accelerate during the climb to stay at the same polar point. Now you further need:

The gradient of air temperature over altitude (lapse rate $\Gamma$)
The local speed of sound $a$, and
The gas constant $R$ of air.

You need to add a correction factor $C$ which has several components: $$C = 1 + \frac{1}{2}\cdot\kappa\cdot R_w\cdot\Gamma_w\cdot Ma^2 + \frac{(1+0.2\cdot Ma^2)^{\frac{\kappa}{\kappa-1}}-1}{(1+0.2\cdot Ma^2)^{\frac{1}{\kappa-1}}}$$

where $\kappa$ is the ratio of the specific heats of air and is 1.405, the index w denotes the wet adiabatic gas constant and lapse rate of air, and $Ma$ is your flight Mach number. $\Gamma$ can vary between -0.004°/m and -0.0097°/m, but if you use the average of -0.0065°/m, this equation can be simplified to: $$C = 1 - 0.13335\cdot Ma^2 + \frac{(1+0.2\cdot Ma^2)^{3.5}-1}{(1+0.2\cdot Ma^2)^{2.5}}$$

which is probably the form you will find in most books which care to cover this topic. The second summand takes care of the reduction of atmospheric temperature with altitude and disappears in the stratosphere, and the third summand covers the additional energy needed for acceleration in terms of flight Mach number.

Acceleration factor over the flight Mach number in the troposphere for the Standard Atmosphere

Now your climb speed becomes $$v_z = \frac{v}{C}\cdot sin\gamma = \frac{v}{C}\cdot\frac{T-D}{m\cdot g}$$

As you can see from the graph above, the correction factor is only important at higher speeds, but will cut the climb speed in half at Mach 2. Some jet aircraft need to climb at a high constant Mach number, and then the aircraft needs to decelerate while climbing. Now the correction factor becomes smaller than unity and the climb speed gets a boost because kinetic energy is converted into potential energy while climbing.

Optimum speeds

To pick the flight speed where the climb speed or the fight path angle reaches a maximum, you now need to describe how thrust will change with flight speed. To simplify things, we can say that thrust changes over speed in proportion to the expression $v^{n_v}$ where $n_v$ is a constant which depends on engine type. Piston aircraft have constant power output, and thrust is inverse with speed over the speed range of acceptable propeller efficiencies, hence $n_v$ becomes -1 for piston aircraft. Turboprops make some use of ram pressure, so they profit a little from flying faster, but not much. Their $n_v$ is -0.8 to -0.6. Turbofans are better in utilizing ram pressure, and their $n_v$ is -0.5 to -0.2. The higher the bypass ratio, the more negative their $n_v$ becomes. Jets (think J-79 or even the old Jumo-004) have approximately constant thrust over speed, at least in subsonic flow. Their $n_v$ is around 0. Positive values of $n_v$ can be found with ramjets - they develop more thrust the faster they move through the air.

The flight speed for maximum rate of climb ($v_y$) is reached at a lift coefficient $c_L$ of $$c_L = -\frac{n_v+1}{2}\cdot\frac{T\cdot\pi\cdot AR\cdot\epsilon}{m\cdot g}\cdot \sqrt{\frac{(n_v+1)^2}{4}\cdot\left(\frac{T\cdot\pi\cdot AR\cdot\epsilon}{m\cdot g}\right)^2 + 3\cdot c_{D0}\cdot\pi\cdot AR\cdot\epsilon}$$

whereas the steepest climb is possible with a $c_L$ of $$c_L = -\frac{n_v}{4}\cdot\frac{T\cdot\pi\cdot AR\cdot\epsilon}{m\cdot g}\cdot \sqrt{\frac{n_v^2}{16}\cdot\left(\frac{T\cdot\pi\cdot AR\cdot\epsilon}{m\cdot g}\right)^2 + c_{D0}\cdot\pi\cdot AR\cdot\epsilon}$$

Nomenclature:
    $c_L \:\:\:$ lift coefficient
    $T \:\:\:\:$  thrust
    $m \:\:\:\:$  aircraft mass
    $g \:\:\:\:\:$  gravity
    $\pi \:\:\:\:\:$  3.14159$\dots$
    $AR \:\:$  aspect ratio of the wing
    $\epsilon \:\:\:\:\:$  the wing's Oswald factor
    $c_{D0} \:$  zero-lift drag coefficient


How to determine the aircraft tail number on a particular flight? How can one determine the registration number (i.e. N123AA) of the aircraft used on a particular flight, having the date, airline, flight number and city pair to start with?

The step-to-step guide:

1. Search for Flight Number:

Get hold of the flight number, such as LH123 or AA789 for instance. This is usually found on your ticket, baggage tag or some email.
Enter it on a site such as www.flightradar24.com or www.planefinder.net.
Note down the registration (such as N1234A or D-ABYT) for the flight on the day your are interested in.

2. Search for this Registration:

Try entering it on for instance www.planespotters.net which is often considered a reliable source for aircraft data.
You should be able to get the aircraft history:

First flight and delivery dates (and hence aircraft age)
Previous operators and registrations.
MSN (manufacturer serial number)


How does ATC handle changing QNH/altimeter values? Obviously for aircraft arriving to or departing from an airport/airspace (as relevant), they listen to the ATIS and/or are given the QNH.

But what happens if that pressure changes? Even a fairly small change can make a big difference when we're dealing with 1000ft of vertical separation (which can be much less by the time we account for human margins for error).

Is the new QNH announced to the entire frequency, or similarly the change of ATIS, which each aircraft would then be expected to check? Or would it just be updated when assigning a new altitude to any aircraft already in the airspace?

In the US, ATC simply broadcasts a notification on whatever frequency is needed. For the terminal environment, section 2-9-2 of the ATC orders says:

  Broadcast on all appropriate frequencies to advise aircraft of a
  change in the ATIS code/message.

The phraseology is usually something like this:

  Attention all aircraft, Bowman information Kilo is now current.

For en route operations (see section 2-7-2) the ARTCC gives altimeter setting updates directly:

  Attention all aircraft, the Nashville altimeter now 2992


Is it possible to fly an EASA registered airplane with an ICAO license? Is it possible to fly an EASA registered airplane with an ICAO license?
My problem is the following: I have obtained my ICAO PPL license in Argentina with 52 flight hours but currently I live in Poland and I'm trying to convert to EASA PPL. I know I need 100 hours for the conversion process but it is not possible for me to travel to Argentina and fly those remaining 48 hours.
Am I able to fly the remaining 48 hours in Poland with my ICAO PPL?

The ICAO doesn't issue licenses, so there's no such thing as an ICAO license. You have a license issues by the Argentine aviation authorities is ICAO compliant, which means it is recognized internationally, and you can fly Argentine registered airplanes anywhere in the world. It does not automatically give you privileges to fly airplanes of any other registration, so no you cannot fly Polish airplanes at this point.
There may be a way to do it, depending on the regulations in both countries. In some cases you can get a temporary license on the basis of your other license, or get some sort of other written permission. You would need to ask the Polish aviation authority what you need to do.


How do you contact ATC if you want to do RC flying in the US? This question talks about flying RC models in US and latest "guidelines" to follow.  One of them is contacting ATC tower if you are within 5 miles of the airport

The seems pretty simple but I'll ask anyway:
How would one go about contacting tower to let them know you are going to be flying RC near airport?

Any specific verbiage that needs to be used?
Can the request be denied?  (Aside from TFRs obviously.)
Does anybody have any actual experience doing this?

Any specific verbiage that needs to be used?

Just tell them you will be flying an RC aircraft with 5 miles of their airport. You should have on hand the location you have in mind. Preferably, know the distance and bearing from the airport, like "3 miles to the southwest". Even better would be a more exact bearing and distance from the airport or a nearby VOR. Also know what max altitude you expect to be at. See the related question for a more thorough description (about NOTAM info, but the same general idea applies): How do I submit a NOTAM?

And if you are wondering how to get in touch with the tower: How can I find telephone numbers for FAA ATC facilities?

  Can the request be denied?

In general I would say yes. One of the main rules of RC flying is to not endanger manned aircraft. If ATC thinks your selected location will break that rule, then you shouldn't fly there. That would be part of why they ask you to contact ATC in the first place.

However, the wording is "provides the airport operator and the airport air traffic control tower...with prior notice". It says nothing about receiving approval, clearance, or permission from them. So I'm not sure if you are breaking any regulations if they say no and you fly anyway, assuming you don't break any of the other rules.

  Does anybody have any actual experience doing this?

Not me at the moment. I will try to update if that changes.


Is the Cirrus Vision SF50 the only civilian monojet currently on production? Since the Piper PA-47 and Diamond D-jet never made it past prototype stage, is the Cirrus Vision the only single engine civilian turbofan jet in production?

What other monojets have made it into production?

Multiple VLJ kitplanes have been on the market for several decades now; probably the most famous of which was the Bede BD-5J which gained fame in the 007 film Octopussy.  BD-5Js flew in airshows with the Coors Light Silver Bullet aerobatic team.
A supersonic VLJ kit called the Bede BD-10 was planned and flyable prototypes were produced.
And multiple companies have come and gone, offering everything from VLJ kits to wishful thoughts such as the ViperJet to MaverickJets to Stratos.
But near as I know the Cirrus SF-50 is the only certified single engine civilian aircraft for sale today.  Eclipse came close with its E-400, but it only remained a technology demonstrator.
The only other single engine jet airplanes on the market are military.  Aero Vodochody's L-29 and L-39 jet trainers are popular with civilians and air race teams as are MiG-15s -17s, -21s and even a few MiG-23s.  Western fighters like the F-86 Sabre, the Folland Gnat, the De Havilland Vampire, and the BAe Hawk are also popular.


Reason for limited engine fuselage on Sukhoi fighter aircraft? To my amateur eye, there appears to be limited fuselage on the engines for Sukhoi fighters from Su-27 on.  That is, the fuselage follows the curvature of the engines closely and there is a definite material difference between the adjacent fuselage.

This seems to be a characteristic specifically of Sukhoi fighters or their clones.

What is the purpose of such a feature?

I realise might be a state secret as to the actual design decision and will accept such an answer but I suspect there is a straight-forward engineering reason.  And perhaps I am mistaken on my assessment.

In a fighter you generally want to optimize fuel and ordinance capacity as well as have competitive performance. This use case generally leads to a design that is different than your average transport category aircraft. The SU-27 has some similar characteristics to the F-14 Tomcat and it should be noted

  The aircraft had a large wing, clipped, with two separate podded
  engines and a twin tail. The ‘tunnel’ between the two engines, as on
  the F-14 Tomcat, acts both as an additional lifting surface and hides
  armament from radar.

Here is a good picture of the engines from the back

(source)

You can see that they could connect the engines but that would increase cross sectional area, disturb the center portion which helps add lift and add weight. If there is nothing they need to stick in there, there is no reason to have such an area on the aircraft. Its also worth noting this is not terribly different from the back of an F-14.

There is not necessarily a whole lot of metal surrounding any aircraft engine.

In the case of the SU-27 it also looks like the engine coverings are unpainted, this may be for heat related reasons.


How are wood and canvas wings able to sustain damp and bad weather? Small aircraft have used and still use at least for wings, and perhaps for fuselage, material like wood and canvas, as visible on the bottom side of this wing:

Robin DR400, source: aeroVFR

Which kind of fabric / which treatment are used? What are the limits:

Are such structure able to withstand bad weather, rain, hail...
Refresh periodicity
Maximum wing load / aircraft size

Wood and canvas are not meant to be used in humid conditions. All outer surfaces are painted, so they are water repellent and can sustain a shower, but when left to the elements, wooden aircraft deteriorate rapidly. The unprotected structure will take up water and will warp. Wet canvas is less taut, many glues lose strength when wet and plywood will come apart over time. When a wooden aircraft has suffered a shower, it should be wiped dry and stored under a roof as soon as practical.

When Junkers introduced all-metal aircraft, they were sometimes the only types to fly because bad weather had made all other types unusable. In the war for independence of the Baltic states during 1919, the rapidly moving frontline made the use of aircraft hangars impossible, and only Junkers airplanes were able to operate on rainy days.

DeHavilland designed a sleek, four-engined airliner, the DH.91 Albatross, in 1936. The last three of seven built were put out of service after a crash in Shannon in 1943. This crash had made the fact obvious that the plywood structure had deteriorated beyond repair in bad weather, despite a fabric-and-dope cover to protect it from the elements.

The good thing about wood is its long-term strength. In stark contrast to aluminium, wood shows no fatigue under stress and allows to design a life-safe structure. When stored and operated properly, a wooden aircraft will last centuries. While the fabric cover will lose strength over time and needs to be replaced maybe every 20 years, the wooden structure needs no rejuvenation.

The maximum size of wood-and-canvas aircraft was demonstrated by the Caproni Ca.60, and that of wooden aircraft by the Hughes H-4. Both were the biggest aircraft of their time (1921 / 1947) and were bigger than a Jumbo Jet if you use wing area to measure size. However, fabric cover puts a low limit on flight speed, and indirectly on wing loading. The DH.91 had only 134 kg/m² and I would be surprised if any wood-and-canvas aircraft with a wing loading above 200 kg/m² exists. Even the plywood-covered DH.98 Mosquito had only 195 kg/m².

Now to the materials: Otto Lilienthal used osier stakes which he covered with nettle cloth. As far as I know, treatment was wax (applied dissolved in a volatile solvent). Early designs used linen or cotton and dope for glueing the cloth to the structure as well as tightening the cover. Since maybe 50 years those natural fibers have been displaced by polyester cloth because it has less of a tendency to rot. Still dope is used for glueing and tightening, and a finish of synthetic resin lacquer is used as topcoat.


Does Ryanair lease any aircraft? Does Ryanair lease any of its aircraft?

And if they do, could the end of the lease period coincide with the start of the low passenger periods? E.g. during the autumn to spring months when holiday traffic is slow, there may be a need to reduce the fleet size.

Ryanair mainly buys their aircraft. Part of their ability to offer cheap flights comes from buying aircraft in bulk and generally only operating a single type of airframe.

They generally buy new Boeing aircraft operating primaraly the 737 series.

  ...had agreed to buy five more Boeing 737-800 aircraft, taking its order
  book to 180 new Boeing planes worth over $16 billion.

Planespotters.net has a nice list of the aircraft in Ryanair's fleet, with the exception of this -700 the rest of the aircraft only show registration to Ryanair as an owner under owner history.


What is this class of bolts called? They can be seen in many —mainly military— jet engines. Do these bolts recibe an special name?

J79: the fasteners are everywhere

EJ200: they can be discerned in the rear flange

12 point flange cap bolt/screw. For high torque applications, less likely to be accidentally rounded-off.
http://www.fbabolt.com.au/12-point-advantages.html


Has anyone tested or flown permeable wings to increase efficiency? See this article: Aerodynamic Performance Improvements Due to Porosity in Wings at Moderate Re
Was wing porosity ever tested/used for real aircraft?
source:

The aerodynamic efficiency parameter clearly shows the positive effect of the permeability for low angles of attack. Having a permeable section on the airfoil surface will provide better aerodynamic characteristics when the angle of attack is less than 10°. This is compatible and in agreement with the observation of other researchers on the effect of the holes on the aerodynamics characteristics of the airfoil that concluded that “at a transitional phase where Re= 40,000 and the angle of attack ranged from 0 to 9°, the L/D ratio was higher compared to the same wing without holes

Not really a great answer but:
Tested - yes I believe so.  Actively used - not to my knowledge/unsure.  The structural complexity of this is challenging, and the cost to benefit probably isn't compelling enough for "traditional" configuration aircraft.
You may find this of interest A History of Suction-Type Laminar-Flow Control with Emphasis on Flight Research or Aerodynamics of Permeable Membrane Wings


Why are the leading edges of the 787's horizontal and vertical stabilizers not painted? The other day I made a question about the engine painting and why it isn't painted at the front part. The answer was: to facilitate anti-ice system.

Source: fervegeon.com

So perhaps the leading edge of the horizontal and vertical stabilizer have a sort of anti-ice system?

And this is not only on the 787; the 777 also has the leading edges unpainted.

Source

And on the 757 and winglet.

Source: flickr

Aircraft stabilizers (both vertical and horizontal) usually have metallic leading edges without paint or anti-erosion paint system. This can be seen in the following figure for Airbus A320 family.

Image from Airbus Structure Training Manual

For aircraft with composite stabilizers, leading edge erosion is arrested by bonding metal to the leading edge, or making the leading edge using metal, as in 787:

  The .. leading edge assembly, ..., with an aluminum edge and a fiberglass and paper-honeycomb skin.

In such cases, (as @Todd points out), the leading edge is usually left unpainted in order to prevent paint erosion. This method is used in the leading edge of the main rotor blades of a number of helicopters. In such cases, either the bare metal is left as is or a protective film is used to combat erosion.


What is the significance of the speed of sound to aerodynamics and aircraft design? Almost every aspect of flight seems to be affected by the speed of sound, somewhat in the same way that nearly everything in physics is determined by the speed of light.

Obviously it's not exactly the same - it is possible to travel faster than the speed of sound, unlike the speed of light - but still, what happens below the speed of sound is quite different from what happens above the speed of sound, while the region close to the speed of sound seems to be a mysterious transition zone.

This isn't just limited to aerodynamics, but also for example to the way the fuel/air mixture burns in a jet engine.

So, why is the speed of sound so significant for aerodynamics and aircraft design?

I presume that it is something to do with a limit on how fast information about the movement (or pressure changes) of air molecules (i.e cause and effect involving air molecules) can travel through a medium of those molecules, and that "speed of sound" is shorthand for that speed.

These movements of air molecules presumably also might be in a form that we wouldn't ever actually hear (frequency too low, frequency too high) or indeed that we wouldn't ever experience as a sound (a single pulse rather than a vibration); in this case "speed of sound" seems to be a rather misleading way of describing it.

I assume that there are analogous issues around the speed of information of pressure changes in water, so that hydrodynamics has to deal with them too. I imagine that because is much less compressible than air, that the speed would be much higher, so perhaps less likely to trouble hydrodynamic designers and engineers.

I think my high-school physics seems to be more or less enough to deal with these concepts, but the language around speed of sound does seem to hide away some of the key ideas, and leads my intuitive grasp in the wrong directions.

Your question shows that you know that the speed of sound is the speed of small pressure changes in a fluid, and moving faster than that will give the fluid ahead no warning of an approaching disturbance. I will, therefore, dig a little deeper in my answer, fully knowing that I can only scratch the surface. I encourage you to ask more specifically if one topic piques your interest but is not sufficiently explained.

Short answer

The speed of sound is significant in aircraft design mainly in:

Splitting the flight regimes into two areas with markedly different flow behavior,
helping to convert a geometry for compressible flow to one which can be calculated by simpler, incompressible methods,
limiting the maximum practical flight speed for efficient air travel and
giving the designer hints how to push this boundary up.

In the combustion of gasses the speed of sound is of less significance than the speed at which a flame advances. This is more complex (PDF!) and depends on the turbulence level, mixing ratio, temperature and the activation energy for the combustion reaction to start.

Explanation

To explain the influence of the speed of sound let me look at the molecular level: Air is a collection of molecules which zip around constantly, bumping into each other frequently. We express this with macroscopic parameters:

The intensity of the collisions is measured as pressure,
the number of molecules per volume is measured as density, and
the speed of the molecular movement is measured as temperature.

At rest, the frequency and intensity of these collisions is the same in all directions. If something moves through the gas at a speed far below the speed of sound, the collisions are still roughly of the same intensity all around. The local intensity may vary, but the molecules will stay equally distributed over the volume. This means that the disturbance will cause a change in pressure, but density will be unaffected.

Similarity methods

Moving faster, the disturbance will cause more intense collisions in the direction of movement and will start to push the molecules together: Now it will also cause weak density changes, and those grow with the square of of its speed. To regain an equal distribution over space, the molecules will now move sideways (and up and down), and the pressure changes will reach farther out to the sides. When you compare the pressure fields at different speeds, the "faster" field will look like one made by a more narrow disturbance at lower speed. This effect is captured by the Prandtl-Glauert factor $(1-Mach)^2$. Scaling the Y- and Z-coordinates of an aircraft geometry (where X is in flight direction) with this factor will result in a geometry which can be calculated with incompressible methods.

At the speed of sound this factor has a singularity, so its application should be restricted to speeds below Mach 0.7. Above Mach 1 a similar transformation can be applied, now using the Ackeret factor $(Mach-1)^2$.

Convergent and divergent flow

The ratio of density changes to pressure changes continues to increase such that in supersonic flow the density changes become dominant. This shows when looking at the flow in an expanding or contracting duct:

Subsonic flow in an expanding duct will slow down to preserve density and will accelerate in a contracting duct.
Supersonic flow in an expanding duct will accelerate further, reducing density and pressure, and decelerate in a contracting duct.

This different behavior is used when accelerating the compressed gasses flowing from an aircraft turbine: First the nozzle diameter is reduced until the gas accelerates to the speed of sound. To accelerate the gas further, now the diameter has to increase again in flow direction. The result is a convergent-divergent nozzle, a typical part of supersonic jet engines.

Wave drag

Now picture the aircraft in slow motion from the side as it slices through the air at supersonic speed: Since pressure increases while the aircraft body expands and decreases once it contracts, you will have two drag components in supersonic flow which do not exist in subsonic flow: Forward facing parts experience higher pressure and rearward-facing parts experience suction. This effect is called wave drag and makes supersonic flight inherently less efficient that subsonic flight.

Transsonic design

The same effect also affects the pressure field around an aircraft which is moving close to the speed of sound. Its movement displaces air and has an effect similar to a contracting duct, accelerating the flow in its vicinity as long as the aircraft grows thicker in flow direction, and decelerating it again over its rear half where it contracts. Once this local flow reaches the speed of sound, it will accelerate further while the aircraft contour contracts, leading to a supersonic pocket of air which will not slow down. It does so eventually, collapsing in a local shock which can cause flow separation and high local drag. Even worse, this separation will affect the whole pressure field and reduce the speed increase, which opens a feedback loop which will make the shock fluctuate. This oscillating shock can excite vibrations, render control surfaces ineffective and can incite flutter.

To reduce the local speed increase the surfaces of airplanes can be swept so the flight speed can be increased without running into the aforementioned transsonic effects. Also, sweep will smear out the sonic effects when accelerating through Mach 1, so the drag peak around Mach 1 can be greatly reduced.

Around Mach 1 it is important to limit the rate at which the aircraft's body expands or contracts over the direction of flight, and the wider-ranging nature of pressure changes allows the designer to expand one part of the aircraft to compensate for the local contraction at a different part when both are at the same position relative to the direction of flight. The consequence of this is the area rule which shows in coke-bottle fuselages or big flap track fairings.


How much energy does unretracted landing gear consume? I am interested in finding energy consumption (% kN) of unretracted landing gear  (on a jet fighter) per hour. Although the shape and size tell us how much drag is being produced I need a general rule of thumb.

While it’s not a true air to air fighter, following are some comparison numbers for the EA-6B Prowler that might help provide a useful ratio for a tactical "fighter type" aircraft.  (Information comes from the NATOPS pocket checklist, under the constant altitude “bingo” fuel divert tables.)
Theses figures are for a flaps-up aircraft, fully loaded with 5 external stores, and both engines operating.  (There are multiple configurations and distances covered in the tables).
Fuel required to go 100 Nautical Miles:

Altitude
Gear Up
Gear Down

Sea Level
3970 lbs
5220 lbs

10,000 ft
3580 lbs
4590 lbs

20,000 ft
3270 lbs
4160 lbs

NOTE:  The gear down figures are worst case and presume a hydraulic failure.  In the event of a hydraulic failure the gear will be blown down pneumatically and the forward main landing gear doors, (that normally re-close to cover the empty wheel well) will remain open.  This imposes some additional drag penalty over and above normal flight with the landing gear down.

You should be able to extrapolate this into something useful, but if you would like more data points please let me know.


How does this RC helicopter keep itself upright? I have this (cheap, beginner-level) RC helicopter:

It has 3 controls: one for climbing and descending (collective throttle), one for yawing (differential throttle), and one for pitching (the tail rotor control). There's no roll control.

This helicopter seems to have a very strong tendency to stay upright. Even if you grab it by the skids in mid-air and tilt it slightly, it will right itself (after first flying in whichever direction you tilted it in).

As shown in the picture, the helicopter has two coaxial main rotors and one tail rotor. The tail rotor is pointed vertically, so that it produces a pitching moment. The lower main rotor is fixed-pitch, but the upper rotor has cyclic pitch controlled by a weighted "balance bar". The balance bar itself is mounted about 45° ahead of the rotor. The balance bar is on a hinge so that the ends can move up and down relative to the shaft. If one end of the balance bar goes up, then the blade closer to it is automatically set to a coarser pitch; meanwhile, as the other end goes down, the blade closer to that end is set to a finer pitch.

It seems very unlikely that this helicopter has any electronic accelerometers or gyroscopes.

So, how does this helicopter keep itself upright? Here's what I can figure out myself:

Suppose that the fuselage accidentally rolls to the right while the balance bar remains upright. Then the rotor's cyclic pitch will be set so that each blade is coarsest when it's in the forward right position, and finest when it's in the rear left position. This will produce a left rolling moment, which will tend to bring the helicopter upright again. (It will also produce an up pitching moment... or maybe a down pitching moment, thanks to phase lag? Or no pitching moment at all? I don't know.)
Suppose that the fuselage and the balance bar both accidentally roll to the right. This will cause the helicopter to fly to the right... which will somehow cause it to right itself? But I don't understand the details of why this will happen.

By the way, I've noticed that the helicopter has a tendency to fly in clockwise circles, especially after being disturbed. (It doesn't yaw during this circular motion; it simply moves in a circle while maintaining a constant heading.) I bet that this tendency is caused by the balance bar somehow, but I don't know how.

(Someone may be tempted to answer, "It rights itself because the rotors are above the center of gravity." That explanation doesn't work, though, because the only way an aircraft can right itself is by means of torque. The rotors will generate this torque somehow, but they won't generate it by virtue of being located above the center of gravity.)

The rotor is gyro stabilized. The balance bar is the gyro. If the machine rolls right, the balance bar wants to stay in a level plane and generates a correction by influencing the rotor blades to go where the balance bar wants to be.

The Bell 2 blade teetering rotor system used on the '47 and the Huey used a much smaller version of the same thing, to provide a little bit of inherent stability to the rotor disc, without inhibiting pilot control.


Does RPM refer to the speed of the engine or the speed of the propeller? I followed discussion here: In a constant speed propeller, why does the RPM drop with a prop pitch increase?, I got confused about the RPM there. As the opening statement said in that discussion, "When a constant speed propeller changes its pitch from low to high, RPM would decrease."

From that discussion said that the RPM there is not the RPM of the engine that used to power the airplane, but the RPM of the propeller itself, how fast it rotating after the pitch was changed. The more pitch of course will produce more drag and as the product, the propeller will rotate slower than previously.

Please you masters of the aviation, please confirm if it is true, the RPM is the rotation of the propeller itself and not the RPM of the engine/turbine used to power of the airplane. That have confused me for quite long time and made me very difficult to understand many explanation.

Yes it is the propeller RPM.  If the engine is direct drive, the engine and propeller RPM are the same.  If the engine has a propeller reduction gearbox, then engine RPM and propeller RPM are different, the propeller RPM being the engine RPM minus the gear reduction.

Propeller RPM is controlled by a flyweight governor, like those spinning ball speed regulators you see on old steam engines.  The flyweights control the oil flow to the prop to vary the pitch. The prop control lever applies a spring preload to the governor flyweights and the amount of preload determines at what RPM there is a state of hydraulic equilibrium so there is no pitch change happening. The speed of the governor is a function of the speed of the propeller in either a direct drive or geared engine.


How many flights can be made under EASA permit to fly? I understand the permit to fly is valid for 14 days, but how many flights can be made in that time?

Assuming you mean a temporary Permit to Fly (rather than a 'permanent' one) for an aircraft normally flying under a C of A, then this site states that:

  You will be required

  to supply a validity period while completing the application.
  Temporary permits are usually issued for less than a month, but you should allow enough time for the flight to be conducted, taking into account weather conditions etc.

I'd guess it is intended for an aircraft to be ferried in order to renew its C of A or at the end of its operational life?


Why are there many pilots who don't say "mayday, mayday, mayday" to declare an emergency? I've been listening to some ATC from some major commercial pilots and in every video where there's been some confusion about declaring an emergency they've used some terminology other than "mayday, mayday, mayday" (here's an example).

Given that how unmistakable saying the word "mayday" is three times in succession to people of different languages and accents, and that it's much quicker than saying "declaring an emergency, declaring an emergency, declaring an emergency", I'm trying to understand why the "mayday" terminology isn't universal. I can't come up with any situation where a triple mayday call won't be the clearest and quickest way to indicate an emergency situation, regardless of whether or not the controllers and pilots speak English or not?

Are there any good reason for not using triple mayday for declaring an emergency? Are there some ATC controllers that wouldn't know what it means? I just don't understand because it seems to be clearly understood every time whereas alternative often aren't, even with two native English speakers.

I'm a controller, not a pilot, so I can only speak from my own perspective.
What we are taught in ATC school is that many pilots are reluctant to use the word mayday because they feel it might escalate a situation unnecessarily and potentially create a lot of paperwork. I guess, mentally, it seems like calling mayday is a significant, irreversible step which can be daunting to take. They might be afraid, assuming the situation goes well, that they have to explain why they decided to escalate the situation, get emergency services involved etc.
That's not really a resonable fear, though. In fact, as a controller, I would much prefer a pilot to call mayday as opposed to trying to make the situation seem less critical than it is. In an emergency, clear communication is absolutely essential. I've had several situations where pilots have had "minor technical issues" or an engine that is "a bit jumpy" or something like that, and it puts me in a really uncomfortable situation because I can't really tell how serious it is. And if I'm in doubt, I will just call the emergency services anyway. On the other hand, the one time I actually had someone call mayday, within 30 seconds, the firetrucks were on the move, all the runways were cleared and the aircraft in question was cleared to land. It's much easier to help you if you are clear about what you need!

Pilots believing themselves to be facing an emergency situation should declare an emergency as soon as possible and cancel it later if the situation allows.
The correct method of communicating this information to ATC is by using the prefix “MAYDAY, MAYDAY, MAYDAY” or “PAN PAN, PAN PAN, PAN PAN” as appropriate. This procedure, which is an international standard, is the single most effective means of alerting the controller to the need to give priority to the message that will follow.

Emergency Communications on SKYbrary
In the end, it is always up to the pilot(s) to decide what is the right thing to do. Every emergency situation is unique, and there isn't (and can't be) rules describing what exactly must be done in every single situation.


How to determine the MAP for an RNAV MDA approach? On this RNAV approach plate, what is the missed approach point for the LNAV MDA approach? How do you arrive at this answer?

The Missed Approach Point (MAPt) is at the runway threshold. The other reference you will see on this RNAV chart is the Visual Descent Point (VDP), at 1.1NM RWY05R, which is at the Minimum Descent Altitude (MDA).

  The concept of VDP was developed by the FAA to encourage pilots to decide to initiate a missed approach prior to reaching the MAP, in a situation where the runway or its environment is not visible at a normal descent angle. Conversely, if the runway is visible at the VDP, the pilot may continue descent, following a standard descent angle to the runway, while being assured terrain and obstacle clearance.[5] The VDP is always located prior to reaching the MAP, and is a more useful checkpoint for making the decision whether to continue on the approach or to go around than the MAP itself.

(Source: WikiPedia)

If the MAPt is intended to be elsewhere, it will be assigned its own identifier.

  "The missed approach waypoint is assigned a pronounceable name when it is not located at the runway threshold."

(Source: AIM 5-4-5.d.5(d))

(Source: www.faa.gov)


How to represent graphically airways and routes? I need to write a piece of software in order to display on a map flight routes and airways. I would therefore like to know if there is a standard way of displaying this kind of information (in particular the kind of icons used for displaying waypoints, departure and arrival airports and so forth).

For example, by Googling a bit I have the chart shown below:

Could someone point me to any additional sources of information that I could use.

I am particularly interested in the icons/shapes to use as well as the kind of relevant information to show. The idea would be that the map is recognizable at once by a professional pilot and provides him all the relevant information.

ICAO Annex 4 contains guidelines for creating aeronautical charts.

The type of information to be shown and required accuracy of information is defined in the different chapters, depending on the type of chart. You will probably be interested in chapter 2 and chapter 7/8.

Symbology and standard colors are explained in appendix 2 and 3.


At what angle-of-attack (sideslip angle) would a symmetric vertical fin plus a deflected rudder have a lift coefficient of exactly zero? For rudder deflection angles of 5, 10, 20, 30, and 45 degrees, at what (negative) angle-of-attack does a fin-rudder combination (including dorsal fin if present) have a lift coefficient of exactly zero?

For the purpose of this question, define angle-of-attack in relation to the chord line of the fin alone, not including the deflected rudder.  If we ignore things like spiral slipstream, deflection of airflow around cabin, etc, it would be the same as the sideslip angle.

Loosely speaking, at this angle-of-attack we could say the force from the deflected rudder is being  exactly cancelled by the force from the fin, though in reality each influences the airflow around the other and the fin-rudder combination acts as a single unit.

To allow for a definitive answer, I'll specify to use the fin (including dorsal fin) and rudder from current -production Cessna 172 S, but feel to post an answer addressing the fin and rudder from any other aircraft.

Note that an aircraft could never sustain a sideslip angle that caused the fin plus deflected rudder to have a lift coefficient of zero, unless some other yaw torque due to something like asymmetric thrust or extreme adverse yaw from deflected ailerons was acting to maintain the slip angle.

Approximations are fine-- obviously an exact answer would involve a complicated computer modelliing project, or actual wind tunnel experiments.  My hope is that existing information relating to the lift curves of symmetrical airfoils with deployed flaps may be adapted to give an approximation of an answer to this question.

Here are two links to Tom Speer's homepage. They show the performance of a couple of flapped airfoils. They aren't symmetrical, but it doesn't matter for what you are doing. What you want to look at is the change in the AoA between adjacent flap deflections.

For instance, it the first link, the there is a series for -3 degrees flap and one for 5 degrees flap. Holding the coefficient of lift constant, you'll see about a 4 degree difference in AoA. Thus deflecting a smallish 20% flap is about the same as rotating the whole foil by half as much.

http://www.tspeer.com/NACA23012/N23012wflap.htm
http://www.tspeer.com/NACA23012/N23012wflap.htm

For a 50% flap, it can be as much as 80% effective. See also Tom's analyses of wingsail performance with various sized flaps at the same site. These are symmetric foils, but you have to do some extrapolation to run the lift down to zero.

Tom worked for Boeing, and has worked on numerous high performance projects including the America's cup wingsails and landyacht projects.


Is AVTUR the same as JP-1 fuel? I read on Wikipedia:

  JP-1 was an early jet fuel specified in 1944 by the United States government (AN-F-32). ... It was also known as avtur.

Is AVTUR the same as JP-1 fuel, and the term "AVTUR" is just a nickname?

AVTUR stands for AViation TURbine Fuel and is equivalent to JP-1, at least in the UK and in as much as JP-1 no longer exists.  I believe that the US military equivalent is JP8 which is Jet-A1 with additional additives, which differs from commercial Jet-A only by a lower freezing point.

Here is the relevant NATO Def Stan (Defense Standard) for Avtur with is the same defstan as for Jet-A1.

This link to an Exxon document gives a great deal of information for various aviation fuel oils.

[EDIT]

From here:

  JP-8 is the military equivalent of Jet A-1 with the addition of
  corrosion inhibitor and anti-icing additives; it meets the
  requirements of the U.S. Military Specification MIL-T-83188D. It is
  the dominant military jet fuel grade for NATO airforces. The UK also
  have a specification for this grade namely DEF STAN 91-87 AVTUR/FSII
  (formerly DERD 2453). NATO Code F-34.


Regulatory Documents To Carry Medical Kits and Defibrillators On-Board Are there any regulatory release documents applicable to Medical Kits and Defibrillators? Question is for Commercial Airlines.
For example for:

Components, you are required to have an EASA Form 1 or FAA 8130-3 release tag.
Standard Parts, you can accept with a CoC (Certificate of Conformity/Conformance).

What are similar standards for Emergency/Medical Equipment?

I can’t comment on ICAO requirements, but for domestic and Flag Operations in the US, §121.801 - §121.805 and §121 Appendix A cover the requirements for emergency medical equipment, including electronic defibrillators, to be onboard scheduled air carrier aircraft and specify training requirements for flight crew members on their use.


Can an aeroplane ever be so far away from an airport that it cannot land? Hypothetically, if an aeroplane with two engines is flying over the ocean and is far away from land, is it not quite likely (but still not very likely at all), that one engine fails and it cannot get to an airport before fuel runs out?

Are there any procedures/precautions that are specifically thought out for this scenario or is it just too unlikely to ever happen?

Can a commercial airplane fly outside the safety area? Yes, there is nothing physically stopping the plane from doing so.

Is it allowed to do so? No, precisely for the event should an engine fail then it will have enough range to get to the airport and make a safe landing.

Extending the allowable range of twin engine aircraft is what ETOPS (Extended-range Twin-engine Operational Performance Standards) is about.

By default commercial airplanes with 2 engines must remain within 60 minutes flying time of an airport in case of engine failure. This was decided at a time where engines were unreliable.

However as engine reliability increased airplanes could have that range increased with an ETOPS certificate.


What determines the "hang point" and control bar position of a hang glider? What is the relation between Center of Pressure (C.P.), Aerodynamic Center (A.C.), "hang point" (the point where the flexible "hang strap" connects to the glider, or the point where the "trike" unit is connected to the wing), and the control bar position relative to pilot's body, if the glider is to fly well, meaning that when I lift hands from bar, the glider will not pitch.

Does hang glider change stability from stable to neutral to unstable when man move his body fore and aft?

I look this video, man must change AoA of wing with bar, so this means no moment(or very little) in this place...
I can't believe how pressure point is so stable/fixed!

What is the relation between Center of Pressure (C.P.), Aerodynamic Center (A.C.), "hang point" (the point where the flexible "hang strap" connects to the glider, or the point where the "trike" unit is connected to the wing), and the control bar position relative to pilot's body, if the glider is to fly well, meaning that when I lift hands from bar, the glider will not pitch.

a) It is important to note that this is a question about trim, not stability.  We are not asking what will happen to the glider after a disturbance.
b) This answer will address the question in terms of Center of Pressure, not Aerodynamic Center, which simplifies things because we don't have to worry about a pitching moment coefficient.  Note however that the Center of Pressure is not fixed.  In general, the Center of Pressure of the aircraft as a whole moves forward as angle-of-attack is decreased, and moves backwards as angle-of-attack is increased.  (Otherwise the glider would be unstable.)
c) We'll assume that the drag on the pilot's body is negligible, i.e. that the "hang strap" hangs straight down when the pilot is not pulling or pushing on the control bar.
d) In the simple case where the hang strap is connected to the glider right at the CG of the glider, or somewhere along a vertical line passing through the CG of the glider, then in order for the glider to be trimmed to fly hands-off, a line drawn through the Center of Pressure, parallel to the Net Aerodynamic Force vector generated by the glider, must pass through the CG of the glider.  Since the Net Aerodynamic Force vector-- the vector sum of Lift and Drag-- must be vertical in a wings-level steady-state glide, this means that the Center of Pressure must lie somewhere on a vertical line passing through the CG of the glider.
e) In the case where the hang strap is connected to the glider somewhat forward or aft of the CG of the glider, then the following must be true: the torque about the CG created by the G-loading on the pilot's body must exactly counteract the torque about the CG created by the Net Aerodynamic Force acting at the Center of Pressure.  If glider's CG is ahead of the Center of Pressure, then the hang strap must be connected to the glider behind the glider's CG.  If the glider's CG is behind the Center of Pressure, then the hang strap must be connected to the glider ahead of the glider's CG.  Alternatively we could tackle the problem by computing the torques around the Center of Pressure rather than around the glider's CG-- we'd get the same result, because we are free to choose any arbitrary pivot point in torque calculation problem.  We'd just have to remember to adjust the force acting at the CG according to the G-loading, in any case where it is other than one.  Note that in the case where the weight of the glider itself is negligible compared to the weight of the pilot-- which is not a bad approximation for some of the lighter "floater" hang gliders-- this reduces down to the simple statement that the hang strap must be connected to the glider somewhere along a vertical line that passes through the glider's Center of Pressure.  In other words the (center of mass of the) pilot's body must be located somewhere along a vertical line that passes through the glider's Center of Pressure.
f) If your meaning of "if the glider is to fly well" is meant to imply not just that the glider is flying at "trim", but also that the glider is statically stable in pitch, then we have the added requirement that if the angle-of-attack increases, the Center of Pressure of the whole aircraft must move aft, and if the angle-of-attack decreases, the Center of Pressure of the whole aircraft must move forward.  This is accomplished largely through washout near the wingtips.  The wingtips, being aft of the CG and having a lower incidence than the wing roots, essentially act like the horizontal tail of a more conventional aircraft.
g) Actually, the above comment is only true when the "hang point" is not far below the CG.  If the hang point is well below the CG-- as was the case in some early hang gliders where the hang point was some ways down on the triangular control frame, well below the "keel"-- then "pendulum stability" could in theory keep the glider stable in pitch even if the "Center of Pressure" were completely fixed as angle-of-attack changed, or even if it moved a little bit in the "wrong" direction-- at least at angles-of-attack not higher than the angle-of-attack that yields best L/D.  Example-- angle-of-attack decreases, L/D decreases, Net Aerodynamic Force vector points further aft relative to flight path and also relative to glider, apparent G-load on pilot (opposite in direction to Net Aerodynamic Force vector) points further forward in glider's reference frame, and this changes the direction of the force exerted by the pilot's body on the glider, acting at the hang point (if pilot is flying hands-off the control bar).  Glider tends to pitch up to a higher angle-of-attack.  Note that we've phrased this description in such a way that we need not assume the G-load is 1G, acting purely vertically.  Note also that this effect is further magnified if the Center of Pressure is high above the CG of the glider itself.  And finally, note that we can reach all the same conclusions, at least in the case where the pilot's hands are free of the control bar, if we ignore the direction of the G-load vector and only consider the direction of the Net Aerodynamic Force vector, which acts at the Center of Pressure, so long as we choose the C.G. of the whole system as the pivot point for our torque calculations, calculating the C.G. of the whole system based on the assumption that the pilot's entire mass is concentrated at the "hang point".

Does hang glider change stability from stable to neutral to unstable when man move his body fore and aft?

Ideally no, at least within the limits of how far the pilot's arms can reach. But in some cases, yes.
Normally in aviation we are concerned with instability when the CG of the system is too far aft.  I'm not aware that this has ever been a problem in actual practice with any flex-wing hang glider.
Some hang gliders can become unstable at high airspeed, so that above a certain airspeed threshold, at progressively higher airspeeds, the pilot must exert progressively less aftwards force on the control bar, or in extreme cases, progressively more forwards force on the control bar, to hold the airspeed constant.  Generally speaking this is caused by flexing of the structure under the aerodynamic load associated with high airspeed.  Certainly if once some threshold airspeed was crossed, the pilot then had to maintain a progressively less forward or further aft body position to hold the airspeed constant at progressively higher airspeeds, that would be clear evidence that the glider was flexing in a way that fundamentally changed its stability characteristics and moved its "Neutral Point".  And this does in fact happen in some cases.
Keep in mind though that "stability", even static stability, is somewhat hard to define in hang gliders.  Do we mean that the glider will return to equilibrium after a disturbance if the pilot exerts a constant force with his muscles on the control bar (including the case where the pilot is exerting zero force), or that the glider will return to equilibrium after a disturbance if the pilot exerts whatever force through his arm muscles is required to lock his body in a constant position relative to the glider (in which case the glider and pilot can be treated as a single rigid body)?  Note that changes in the L/D ratio, which change the direction of the net aerodynamic force generated by the glider, and thus the direction of the G-loading acting on the pilot's body, will change how far forward or aft the pilot's body tends to "hang" relative to the control bar when the pilot exerts no muscle force, so these two definitions of stability clearly are not the same.  In some cases involving rather steep dive angles, as angle-of-attack is progressively decreased and airspeed is progressively increased, the pilot may experience a need to exert a progressively less aftwards force or in some cases even progressively stronger forward force (the opposite direction that we really would desire to be the case) on the control bar to keep the glider's airspeed constant, yet the pilot's body position is actually being moved progressively further aft (which is equivalent to moving an aircraft's control stick further forward, and is intuitively consistent with our concept of a stable of aircraft.)  A hang glider pilot would typically describe this as "a loss of normal bar pressure", yet the relationship between bar position and airspeed may remain "normal", with the bar having to be placed further aft to reach higher airspeeds.
It's complicated-- some force vector diagrams are really needed to make the point completely clear.  But in short, the relationship between "control stick position" and "control stick pressure" is less straightforward in hang gliders than in most "conventional" aircraft (with unboosted mechanical control linkages).


Can planes fly without wings but with a horizontal stabiliser? There are X-Wings in Star Wars and I have been making a model recently. So can planes fly without wings but with a horizontal stabiliser?

This is a question about terminology more than aerodynamics. All horizontal surfaces can contribute lift, and also the fuselage does. Take missiles, for example. They create dynamic lift like an aircraft, but only with their fuselage and fins by inclining both against the oncoming flow. Speed compensates for the small area and poor lift efficiency of slender bodies. Since a fuselage is unstable in pitch and yaw, you need to either add fins or something else which creates perpendicular forces at the far end of the fuselage to control it. Like missiles do.

AIM-9L (picture source)

An X-wing will make a horrible model aircraft. But with enough thrust even this design will fly.


What is the difference between a service manual and a maintenance manual? What is the difference between service manual and maintenance manual? Because sometimes I found one, sometimes I found another, the content is similar but not the same. What are the differences and purpose of each document?

I don't see the definition in the FAA site.  But based upon experience in military aviation we look at them as synonymous terms. The terms differ at the frequency and level of care provided by maintenance technicians.
https://www.faa.gov/guidance/
Maintenance manuals are aligned with the term Operations & Maintenance manual, which is for daily operations maintenance and corrective actions.
Service manuals are for longer interval maintenance practices when the aircraft is in the hangar undergoing longer maintenance projects.
Ex:  engine flush vs engine disassembly


What color are cockpit weapons control? Forgive my ignorance in my wording of this question.

I assume the weapons control in a military aircraft cockpit are colored differently from other instruments the pilots has to look at considering U.S. military pilots are not allowed to be colorblind, Why do some Air Forces not allow Color Blind Pilots?

If that is correct, what color are the weapons controls?

Please add any appropriate tags I missed.

Depends what you define as a bonafide control? Lets look at the F/A-18C

Releasing a weapon from the aircraft is not just a case of pushing the red weapons release button or trigger. A MASTER ARM switch is included to prevent the accidental discharge of weapons, (and in some other aircraft countermeasures and stores jettison functionality). The MASTER ARM switch is not painted.

However this is still not enough, as the systems of the aircraft have to be programmed to release the selected weapon, perhaps by selecting the relevant weapon as configured in the STORES MFD page. In (most?) Hornet (blocks?), the MFD symbology is monochrome green regardless of the page selected.

There are shortcuts into weapons master modes though: The lowest hat switch on the stick is a shortcut into A/A master mode, with each direction selecting a different A/A weapon and the MFDs being changed to relevant pre-set pages. This switch is black, much like the other stick controls.

So, the pilot, colour-blind or not, has to be able to locate multiple controls in order to release anything from the aircraft, which are neither painted bright red, the same colour or any differentiating colour from their surroundings, except for the weapons release itself.

There are enhanced answers to be provided regarding workflows, muscle memory, the actual shape of the controls, training and the priorities of selection. There are much better qualified people here to talk about those.

P.S. I would provide an image to show where these controls are, but I'm not able to currently.


Does a turboshaft's exhaust provide any thrust? See the below figure. Does the exhaust gas coming out through the nozzle have any force?

If the exhaust gas has force then how is it controlled by the helicopter?

Yes, the exhaust of any engine does provide some thrust (very little for piston engines).

The amount of jet thrust for turboprops is in the range of 4%-15% of propeller-produced thrust. For a helicopter, it's only 2%-4% of rotor lift, but it's directed backwards. The helicopter in your picture would get 100-150 lbf of thrust from the engine exhaust, compared to 3,000-5,000 lbf of lift provided by the rotor.

This thrust is not controlled in any way. It helps a little in forward flight, and is compensated for by cyclic pitch in hover or maneuvering.

Residual thrust can be calculated, but is rarely specified in most sources. You can find some analyses of residual thrust for common engines online, such as for the PW-100 series, or for the Allison T-56.

The T-56's residual thrust, per above analysis, is 750 lbf, with 3442 kW on the shaft. That allows the T-56 engine to produce up to 10,000 lbf of thrust when used as a turboprop with a propeller. If used in a helicopter, such an engine would be able to deliver 25,000-40,000 lbf of lift with a rotor.


How many LRUs are there in modern aircraft like a Boeing 747 or an Airbus A380? I work in the Avionics industry. I always wonder, how many Line-Replaceable Units (LRUs) are present in modern aircraft like A380 and Boeing 747.
I understand that number of LRUs varies from customer to customer. I just want to know in general how many LRUs present in an Aircraft.

It's difficult to precisely answer your question, because they are not public information and depend on the aircraft sub-model, customer configuration, etc.
I built this approximate plot using the data from [1] that shows the trend in Airbus models:

It should be quite surprisingly to you that the A380 has the same number of computing units of the A340. This is because of the transition from the LRUs design to the IMA design (see this question for the difference), that allows to considerably reduce the number of computing units required, called Line Replacement Modules (LRMs). Airbus claims they were able to cut the 50% of the required computers in A350 thanks to IMA approach [2]. A similar claim has been made by Boeing for the 787.
The data does not show the recent aircraft models, but I expect they are still around 100 or slightly above on newer aircraft like A350XWB (which shares most of the A380 avionics). This is because increasing the number of avionics computers is very problematic as it has terrible cascade effects:

This conceptual diagram may look exaggerated, but consider that Boeing claims the reduction of 2000 lbs (~900kg) switching from the LRU approach to the IMA approach [3]. Saving 900 kgs means adding 10 people to the cabin, and it's easy to imagine the money saved by the airlines in the lifetime of the aircraft.
Anyways, check this video which shows the avionics bay of A350XWB: pretty impressive.


Are the audible warning standardized? This question focus on airliners but is not restricted to Airbus and Boeing.

In aircraft, there are a lot of audible warning (usually working along with visible warning). This include auto-pilot disengaged, stall, GPWS, wrong flaps configuration, low energy, ... (I'm not aware of all audible warning that are implemented)

Given one of this warning, is the sound associated to this warning the same in different aircraft? For example, does a GPWS warning sound the same in Boeing, Airbus, Embraer,...?

General alerting sounds are chosen by the manufacturer of the cockpit systems. Stall warnings, autopilot disconnect alerts, fire alarms etc. are all not standardised, with the exception of the words spoken by the GPWS/TAW and TCAS system.

Whilst the exact sounds of these two systems are not standardised in the sense that a standard .wav file is to used by all manufacturers, the phraseology in their aural alerts is standard.

The Ground Proximity Warning System (GPWS)/ Terrain Awareness Warning System (TAWS) will have to issue an aural warning 'PULL UP, PULL UP' if the aircraft is approaching terrain, and 'GLIDESLOPE' when the aircraft is deviating far below the glideslope. Also the cautions ('SINKRATE.', 'TERRAIN.', 'DON'T SINK.', 'TOO LOW, GEAR.', 'TOO LOW, FLAPS.', 'TOO LOW, TERRAIN.' and 'FIVE HUNDRED.')1 are standardised. Whether it is a male, female or robotic voice is up to the manufacturer of the system, but the words are uniform across all systems. Additional cautions and aural alerts are allowed as long as they are clear and unambiguous.

The same holds for TCAS; 'TRAFFIC, TRAFFIC.', 'CLIMB, CLIMB.', 'DESCEND, DESCEND.', 'INCREASE CLIMB,', 'INCREASE DESCENT,', 'REDUCE CLIMB,', 'REDUCE DESCENT.', 'CLIMB, CLIMB NOW.', 'MAINTAIN VERTICAL SPEED, MAINTAIN.', 'LEVEL OFF, LEVEL OFF.' (supersedes "ADJUST VERTICAL SPEED, ADJUST), 'MONITOR VERTICAL SPEED.', 'CROSSING.' and 'CLEAR OF CONFLICT' are all standard phrases which have to be implemented in this exact wording(2).

1FAA TSO C151c Table 4-1 (pages 16 -- 18)

2 RTCA DO-185B, Minimum Operational Performance Standards for
Traffic Alert and Collision Avoidance System II (TCAS II)
Version 7.1, Volume I section, 2.2.6.3: Aural annunciations)


Are there any situations where it's beneficial to disable altitude reporting in TCAS? Having altitude reporting enabled persistently seems like a pretty common-sense thing to do -- it enables RA in addition to TA, the benefits of which are fairly obvious. That said, most of the common transponders installed in GA and commercial aircraft have either the option to explicitly turn on (typically on GA) or to explicitly turn off (typically on commercial) altitude reporting.

Are there any normal or abnormal flight conditions which would make it impractical, unsafe, or impossible to have altitude reporting turned off?

The only time it would make sense as a normal or abnormal procedure to turn off altitude reporting is when your static encoder system is reporting no or incorrect altitude information. Especially in today’s world of mandatory Mode C and/or ADS-b in certain airspaces. ATC in those airspaces may have you cycle your system temporarily. Or, they could have you permanently  turn it off while you either land as soon as practicable or vacate the airspace until the system is repaired.


When does a 'flaps/slats misconfiguration' warning sound during takeoff roll? If a flight crew begins their takeoff roll and don't have slats and flaps configured properly, when does the system know when to sound the alarm? If it actually sounds during the roll itself, that seems a bit too late, as there then isn't enough time to deploy those surfaces. And it doesn't seem likely that the crew would have the time to look up a new V2 to figure out at what speed they can take off without flaps and slats.

For the Boeing 727-100 and 747-100/200 the answers are:

  when does the system know when to sound the alarm?

When the thrust levers are advanced to begin the takeoff, if flaps and slats and a few other things are not in takeoff position, a loud horn sounds.

  figure out at what speed they can take off without flaps and slats.

Takeoff is not permitted without flaps and slats. You have to have them. Depending on how many were not in the takeoff position, the takeoff speed would be inordinately high, require runway lengths not available, and might be above the maximum tire speed. In short, as I said, you have to have them.

On the aircraft I mentioned, if you're starting the takeoff roll from a stationary position and advance the thrust levers at the usual rate, you'll get the horn before your ground speed is little more than walking speed, if that. The thrust lever position, while above anything you would use for taxi, is not anywhere near what you could take off with. It's not far along the travel of the thrust levers from idle to full throw.

The procedure taught for a takeoff warning horn at the carriers I flew for was to bring the thrust levers all the way back to idle, brake to a stop, and request clearance from the tower to vacate the runway. Once off the runway, brake to a stop and troubleshoot the problem. In no case were you to troubleshoot the problem while moving on the runway.

In practice, if there was no traffic, once stopped some would stay on the runway to troubleshoot after getting tower clearance to do so. The assumption was that the problem would be quickly corrected and takeoff resumed.

Troubleshooting was a simple matter of checking that everything that was supposed to be done had been done. For the 747, the items that would trigger the takeoff warning horn were:

leading edge flaps not extended, there were lights on the f.e. panel that showed their individual positions,
stabilizer not in the green band, there was an indicator by the trim wheel on the center console.
flaps not in 10 or 20 degrees position
body gear not centered, lights on the f.e. panel showed its position.
speed brake lever not in the down detent.

Recalling these items from memory was a standard question on oral exams.

If everything had been done that was supposed to be and you still got the horn, it was usual to recycle everything. If that didn't fix the problem, which if the problem was the warning system itself would be the case, it was back to the gate for maintenance.


If all engines fail in a Boeing, how heavy are the flight controls? Consider a large airliner with no Fly-By-Wire controls, such as a Boeing 747 or 767. All the engines have failed, the Ram Air Turbine is deployed, and the plane is gliding.

How heavy are the flight controls in this situation?

What about Boeing 777 and 787, which are Fly-By-Wire aircrafts?

EDIT: Regarding the question "how heavy", I would like to get a number, or if that's not possible, a description from a pilot whether it is physically exhausting to hold the controls.

A 747 doesn't have a ram air turbine (RAT), but it has an APU that can power the hydraulics. If that can't be started, a 747 can still function using the hydraulic pressure from each engine's windmilling. Computer based training video here.

Trivia: if all hydraulics fail, but engines are running, a 747 is a dead stick, but can be steered using the engines. A complete loss of engines and hydraulics is never in the emergency procedures of a heavy airliner.

In a 767, assuming the APU also won't start, the RAT will provide hydraulic pressure. Controls will be heavier (but not impossible) the slower the plane is.

Source: 767 QRH, 13.14

767 RAT speed below 20,000 feet is >200 IAS, any slower, hydraulic pressure will be lower.

A 787's artificial feel system (Elevator Variable Feel) will revert to direct mode, feel will be based on flap position, and during approach feel forces will be reduced to make the landing easier.

Source: 787 FCOM, 9.20.12

Trivia: a 787's RAT can supply 5000 psi of hydraulic pressure.

In a 737, the crew will use manual reversion mode, for roll and pitch it's control tabs, like an MD-80, heavy but not impossible.

For all types, it will be almost impossible to find data on the exact amount of force needed. Also for all types, if no APU or RAT, windmilling can still provide some hydraulic pressure.

Same for the pilot feedback, as very few flew all-engine-out heavies. And adrenaline rush would almost certainly skew their post-flight judgement.


Why are aircraft windshields not unbreakable? From a materials engineering standpoint, cockpit windshields,
  typically weighing between 25-40 pounds, depending on the type of
  windshield build, are typically constructed of several layers:

  (1) A glass face-plate, roughly 1/10" thick;

  (2) A Polyvinyl Butryal (PVB) layer, about 1/8" thick;

  (3) A stretched acrylic layer, approximately 1" thick;

  (4) An additional PVB layer, nearly 1/10" thick;

  (5) An additional stretched acrylic layer, approximately 1" thick. Source

But even then, it is not unbreakable. We have had several cases of broken windshields, as mentioned in above source. Also, the co-pilot  in Varig Flight 254 crash was wounded from a tree, which struck his shoulder after crash landing.

So, why are aircraft windshields not made from unbreakable Glass Ceramics ?

Windshields

In the US, aircraft windshields must conform to 14 CFR 25.775 - Windshields and windows

  (b) Windshield panes directly in front of the pilots in the normal conduct of their duties, and the supporting structures for these panes, must withstand, without penetration, the impact of a four-pound bird when the velocity of the airplane (relative to the bird along the airplane's flight path) is equal to the value of V C, at sea level, selected under § 25.335(a).

So, windshields are not indestructible partly because they are not required to be.

I guess windshields are not required to be indestructible because bodies like the FAA probably take the view that when an aircraft impacts a massive object at high speed, say a large mountain at 500 MPH, it is of no consequence whether the windshield survives intact at the bottom of a large crater of tiny aircraft fragments. It serves little purpose to make the windshield substantially stronger than the surrounding airframe.

Glass-Ceramics

Glass Ceramics are not unbreakable. For example a glass-ceramic might have a bending strength of 340 MPa (Ref p9)