Why were the space shuttle’s flight controls only certified for low airspeeds? Somewhat ironically for a vehicle designed to reach orbital speed, the space shuttle’s flight control system was only certified for airspeeds up to 333 KIAS, and exceeding 470 KIAS (the maximum allowable in dire emergency) would have led to an immediate loss of control.  This constraint was the reason for at least two, and possibly three, of the low-survivability zones in the event of a triple engine failure (go to page 14 of the linked PDF and look at the number of times it says “excessive EAS during pullout”1).

In contrast, essentially all modern jets have flight controls certified to well over 500 KIAS, and capable of safely controlling the aircraft at airspeeds considerably higher still.

Given the lethal effects of an overspeed-induced loss of control during an emergency situation, why were the shuttle’s flight controls certified with such a ridiculously low maximum airspeed?

1: Technically, it uses KEAS (knots equivalent airspeed) instead of KIAS (knots indicated airspeed), but, assuming that your airspeed indicator is properly calibrated and the pitot-static system on your aircraft is working properly, KEAS=KIAS.

From what I understand the limits seem to be related to contingency maneuvers such as RTLS or TAL where the external tank was still attached. These were only possible after the solid rocket boosters had been jettisoned, but while the external tank was still attached.

I couldn't find any altitude references for RTLS, but according to this
TAL required the shuttle to fly to around 350.000 feet. RTLS would have been at a lower altitude and required a 180 degree turnaround. Doing these maneuvers with the ET places a higher stress on the vehicle resulting in black zones. Wayne Hale also mentions this in an old blog post, but doesn't go into specifics.


What happens if the pilot tries to do a loop in a full-motion simulator? I know that full-motion simulators, typically used to train pilots on commercial aircraft, have limits to how much they can tilt. They can't go upside down, for example. What happens if the pilot goes outside those limits, by doing a loop for instance? Does the simulation continue with the loop and the just the physical tilt effect gets ignored, or does the whole simulation shutdown and reset?

The motion systems in simulators do NOT mimic the airplane's attitude. Whether the airplane is right-way-up, upside-down or inside-out has no effect on the motion system.

The motion system is used to move local gravity around to mimic the accelerations felt while flying. eg. To mimic the braking during landing, the cab tilts forward, moving local gravity forward relative to the cab, making it feel like you're hanging in the straps due to the braking. Note that when you are braking during landing, the cockpit is level, and yet you are still propelled forward.


How to recognise the Tupolev Tu-154, Tu-204, and Tu-334? Tupolev used to be the one of the largest airliner manufacturer comparable to Boeing and Airbus. However since the dissolution of USSR, its airliner business is declining, and only operated by a few ex-USSR or communist countries' airway. Currently narrowbody Tu154, Tu204 and regional Jet Tu-334 are still active.

How to recognise Tu154, Tu204 and Tu-334, compared to their rivals with similar size?

First, the Russian airplane industry is organized differently from its Western counterpart. While both have research organizations and aircraft manufacturers, in Russia and Ukraine design offices are between both. In Russia, TsAGI does most basic research and develops aerodynamic designs which are then put into fully developed aircraft by the design offices (Antonov in Ukraine, Ilyushin, Tupolev, Mikoyan (MiG), Sukhoi, Yakovlev or Beriev in Russia). The finished designs are then transferred to manufacturing plants like Aviakor in Samara which churn out the actual hardware. If the demand for Tupolevs goes down, Aviakor will switch to Antonovs or Ilyushins, so the fluctuation for products of a particular design office can be evened out.

One consequence of this arrangement is that aerodynamicists have the first say on the shape of the aircraft. Russian designs generally have more refined aerodynamics, at the cost of a higher manufacturing effort.

Older airliners from the Fifties and Sixties were designed with dual use in mind, and many of them had a glazed nose, which makes them easy to distinguish from their contemporaries from the West.

Tu-154

This design from the 1960s is similar to the Hawker-Siddeley Trident or the Boeing 727. Its production surpassed 1000 aircraft and ended only in 2013.

Tu-154 of Uzbekistan Airways (Picture source).

Four items make it easy to identify:

The landing gear uses six-wheel bogies, which is unique among three-engined airliners.
For aerodynamic refinement, heavy use of Küchemann bodies is made in the T-tail intersection and for stowing the landing gear.
The aerodynamic refinement also shows in the length of the conical section of the fuselage, which takes much longer to reach its cylindrical shape. This is a consequence of putting aerodynamics before ease of manufacturing.
The central air intake is of elongated shape, while that of the Boeing 727 is round and the one of the Trident was a heavily rounded triangle.

Tu-204

Intended to replace the Tu-154, this two-engined airliner flew first in 1989 and looks similar to a Boeing 757. It is still in production, but less than 100 have been produced.

The main differences from the Boeing 757 are:

Since the cockpit of the Boeing 757 is identical to that of the Boeing 767, it looks oversized for the narrow-body fuselage of the 757. In comparison, the cockpit windows of the Tu-204 are smaller and look more natural.
The top of the 757 fuselage slopes down over the full length of the vertical tail, whereas the downslope of the Tu-204 fuselage starts only at the rudder.
To distinguish the Tu-204 from an Airbus A320, look at the wheels (four-wheel bogey on the Tupolev, two-wheel gear on the Airbus) or the APU exhaust, which in case of the Tupolev is a small tube sticking slightly out from the fuselage, while the Airbus has a metal-colored cover at the end of the fuselage.
Another difference is the Tu-204's much longer fairing for the wing-fuselage intersection.
Also, the winglet shape is quite different between those three types.

Tu-334

The Tu-334 is a shortened Tu-204 with fuselage-mounted engines. Only two have been built, so the chances to spot it are remote. The new location of the engines allows it to use a very short landing gear, and the shortened fuselage gives it a rather stubby, unique appearance.

Similar airplanes are

Short versions of the DC-9 like the DC-9-10 or the DC-9-15; here the low bypass ratio of the DC-9 engines results in much smaller intakes and longer engine nacelles. Also, the fuselage of the DC-9 is more pointed ahead of the cockpit. The cockpit windows of the DC-9 have a separate window above the regular row of windows for better visibility in steep turns, which is missing in the Tu-334.
Comac/Avic ARJ-21, a Chinese development which is based on the DC-9 fuselage, but has a modern, Antonov-developed wing and high-bypass-ratio engines. Here the upper cockpit windows are missing, so the best way to tell the Tu-334 apart from the ARJ-21 is again the more pointed fuselage nose of the ARJ-21, the bigger fairing of the ARJ-21's wing-fuselage intersection, the lower aspect ratio of the Tu-334 vertical tail and the rear end of the fuselage, which ends with a vertical edge in case of the ARJ-21.


Would airline got charge for late on depart from the gate due to passenger boarding issue? If so, what is the charge like? Airline wants the boarding be completed as quick as possible.  I wonder what is the cost for any delay? On the other hand, I want get a dollar value for any delay and see the justification value for any improvement.  So, I wonder airline would get additional charge for the gate on delay ?
Thanks.

Chain of delays

Being late on departure is likely to cause you to be late on arrival, this has the following immediate consequences:

you may be delayed even more on arrival in congested airports (e.g. London) if you miss your scheduled landing timeslot, being forced to circle around for, say, 20-30 minutes wasting fuel and time until you're allowed to land; I have been on a flight like that.
the main immediate effect is that some transferring passengers may be delayed because of this, which may possibly result in costs to you to reroute or compensate them.
you likely have scheduled the same airplane for the next flight with a minimal turnaround time; and delayed arrival may easily cause your next flight of that same airplane to be delayed (again, I have personally seen delays caused by this)
you likely have scheduled the same crew for the next flight, and in certain situations that delay may push them over the allowed work/rest schedule, so they are prohibited to fly their next (last in their shift) flight - so you need to rapidly find a new crew for the scheduled flight or cancel it.

For these reasons, a single delay sometimes might cause a chain of other delays or even cancellations, and become quite expensive.


Why have different departure hall and arrival hall locations? For most of the newly built airport terminals I've seen, the arrival halls are located on the lower level and the departure halls are on the upper level.

What is the benefit and disadvantage when:

Departure halls are above arrivals.
Arrival halls are above departures.
Arrivals and departures are on same level but in different places.

How do arrival/departure locations inside the airport impact the placement of airliners at the gates?

I'm not sure if 'most' airports have this design. I've seen both cases (i.e. departures below arrivals and the other way around). Atleast for small airports, a few reasons come to mind for having departures above arrivals:

If arrivals is on ground floor, the baggage can come easily from the aircraft. This reduces complexity, especially in small airports.
Most of the aircraft are boarded using jet bridge- makes sense if they are boarded from first floor (or above); of course, jet bridges can also be used for disembarking, so this is questionable.
The upper floor offers a good view to enjoy while shopping or eating. This maybe the reason as more time is spent in departure. Also, if the arrival is in ground floor, one can simply 'walk out' as no checks are there (international flights are a different case).

None of these arguments are convincing, I guess.


What is this long Bar/Rod on the Boeing MH-47G? This MH-47G has a thick, long rod at the bottom left front. What's the purpose of this thing?

This is the In flight refueling probe, which can be used for refueling while the helicopter is airborne:

(image source: pinterest.com)

It is almost 9 meters long to make sure the rotors do not hit the refueling drogue. You can see it in action in the following picture:

(image source: chinook-helicopter.com)


What if I fail to receive FAA Medical Certificate via mail? My AME just called to let me know that my FAA Class 3 medical certificate had been issued and was sent to me via airmail from United States to China. Since April 7th, 2014 when it was dispatched, it has been over 2 months now and I still don't receive the certificate.

I called local postal office and was told I needed to find a 13-digits tracking number (e.g. AB123456789CD, 13 digits, with 2 alphabetic letters at the beginning and end of the string and 9 digits in between) of the mail so they could check for me. The problem is I don't have it and I have no idea where to find it on the https://medxpress.faa.gov website.

Can anyone help me on how I can get this number? What are my options should I fail to get ahold of the certificate?

You're really asking a postal service question (and from my experience with the postal service you're probably out of luck, even if you had the tracking number: Unfortunately when the USPS hands packages off to other nations' postal services that's often the last tracking update - either the item gets to its recipient or not, but it's gone into a black hole).

In any case the answer to the aviation/FAA half of the question ("I haven't gotten my new medical certificate yet - what do I do now?") is much simpler, and involves more helpful people than international mail: "Contact the Medical Certification branch" (see here).

In this particular case you want to actually call them on the phone and speak to a human being to determine (a) if they have a tracking number, and (b) how to proceed to ensure that you have a valid medical certificate in your possession.
I know in cases like this the Medical Certification branch can also fax you a copy of the certificate (I believe these are valid for 60 days) while you try to get things straightened out and get the actual paper certificate in your hands.


What would happen if I bank a helicopter then pull hard on the cyclic? Prompted by Simon's response to my question here, I will ask another question about reality.

If I:

Start with a AH-6 helicopter in a hover
Pitch down slightly to accelerate, then slowly increase collective to gain speed until max collective is applied and the airspeed is near Vne
Roll the helicopter to about 70 degrees of bank
Pull back hard on the cyclic

(Basically it can be described as trying to fly a Red Bull race in a helicopter)

What would happen?

I can't provide references specific to this manoeuvre. The main texts on helicopter dynamics only generalise since such a manoeuvre isn't flown in real life or tested during certifications.

When you raise the collective, you increase the pitch of the all of the blades together.  In most flight regimes, this means increased angle of attack and therefore lift.

The cost of producing lift is drag which always increases.  So, there is a direct relationship between the collective lever and power required from the engine.

As you continue to raise the collective, AoA and lift will increase until you reach a point where the power required to overcome drag exceeds the rated power of the engine.  Most helicopters permit operation beyond the maximum for a short time, e.g. 2 or 5 minutes, typically during take-off.  Continue to increase AoA beyond this limit and drag will exceed power available and the rotor RPM will drop (a reduction in rotor RPM caused by drag exceeding power through too much collective is known as "droop").

As RRPM decreases, so does lift and the helicopter will descend.  The relative airflow into the disk now has a greater horizontal component and the angle of attack will increase, either causing yet more drag or stalling the rotor.

So there is some point, before the collective is pulled fully up where maximum lift is attained.  Beyond that, lift reduces until recovery or a stall.  The only time I have ever raised the collective fully is during "full and free" movement checks with the engine off as part of pre-flight.

Your hypothetical manoeuvre:

From VNE in level flight, you roll into a 70 degree bank.

The TRT( total rotor thrust) is always acting perpendicular to the plane of the rotor disk.  Since you are in a bank, and I assume it is at constant altitude, there is a large horizontal component of the TRT to maintain the acceleration in the turn and a large vertical component to oppose gravity and maintain altitude.

This requires a lot of power which means a lot of collective and a lot of drag.  I don't know for the AH-6, maybe there's a chart relating bank angle and speed to power but let's just say, your collective is going to be somewhere near the max lift point with a manoeuvre like this.

Since you are now accelerating with max power, your speed will decrease.  Therefore lift will decrease and since you are at or near max power, you will begin to descend.

So far, this is probably recoverable simply by rolling level and reducing collective until back inside the performance envelope.

The second part is where it definitely will go wrong. In all helicopters with a non-rigid head, including the AH-6, the disk will adopt a new plane of rotation rapidly.  There is then a significant delay before the fuselage adopts a new attitude which puts the TRT perpendicular with the aircraft Z axis.

Pulling hard on the cyclic will very rapidly put the disk into a new plane of rotation which has a large angle between the plane of rotation and the fuselage.  Additionally, the disk loading will increase dramatically and the blades will be forced downwards.  On an articulated head such as that on the AH-6, droop stops will prevent the blades from bending down so much that they fold.

Hitting the droop stops a little is uncomfortable with a lot of vibration. Pulling hard on the cyclic at VNE in a 70 degree bank at or near max power may cause failure of the droop stops or blades or cause the static droop stops to contact the  mast possibly leading to rotor separation.

Let's assume that there is no failure.  The plane of rotation will be at the maximum permitted by the physical limits of the flapping hinges which is greater than the angle between the plane of rotation and the fuselage.

Therefore, I would expect a hard pull on the cyclic to result in a blade hitting the tail or the top of the cockpit.  Either results in destruction of the aircraft, either as it breaks up at that point or when it hits the ground 10 seconds later.


What are some Anti-icing solutions for small planes without an on-board system? When flying during the winter, is there anything I can do to increase my margin of safety, such as anti-icing fluid?

I've heard of people spraying their aircraft with propylene glycol, but am curious to know if this will do anything to prevent ice after takeoff, especially without a thickener.

There are many things you can do to mitigate the risk of icing, but most involve prevention via avoidance of icing conditions.  For example, getting a good weather brief, knowing where the freezing level is, staying away from visible moisture, and finally, recognition and taking evasive action.

Pre-flight deicing liquids are not intended to prevent ice accumulation in flight.  As you noted, they will just blow off in the slip stream.  Anti-icing fluids such as TKS are very effective, but the aircraft must be equipped with a system to properly disperse it to be certified for flight into known icing conditions.

If you are not instrument rated it gets easier, because if you stay out of clouds you won’t ice up.  But if you want to fly IFR in the winter, (or any time you could be IFR in the  temperature zone where ice could form) your best method of increasing the safety margin is to arm yourself with knowledge, practice operational risk management, and exercise sound judgment to keep yourself out of icing.


What additional airspace will be off-limits to an a/c w/ no elect. system or xponder under 91.225 (ADS-B) compared to under 91.215? Under FAR 91.225 (requirement for ADS-B-out), what additional airspace over the 48 contiguous states in the U.S. will be off-limits to an aircraft w/ no electrical system or transponder, compared to the current situation under FAR 91.215 (requirement for Mode C transponder)?

The key difference appears to be that under the new regulations, aircraft without engine-driven electrical systems and without ADS-B-out will be barred from flying above any underlying Class B or Class C airspace, AT ANY ALTITUDE.  This is a function of FAR 91.225(e)(2).

Currently, aircraft without electric-engine-driven electrical systems and without transponders are barred from flying above underlying Class B or Class C airspace only up to 10,000' MSL.

Also, under the new rules aircraft with no engine-driven electrical systems and no ADS-B-out won't be able to fly above 10,000' MSL under a Class B or C shelf. Again this is the result of FAR 91.225(e)(2).  The only such airspace appears to lie under some outlying Class B airspace near KSLC.

This answer is based on an understanding that the restrictions imposed by FAR 91.225(e)(2) are only construed to apply within the lateral boundaries of (i.e. above, below, or within) Class B or Class C airspace.  This understanding of FAR 91.225(e)(2) is congruent with the way the FAA is currently interpreting the related regulation FAR 91.215(b)(3)(ii).  Other interpretations of FAR 91.225(e)(2) could produce much more restrictive results for aircraft without engine-driven electrical systems and without ADS-B-out, but are problematic for other reasons, and seem inconsistent with the intent of the regulation.

For more on why this seems to be the best interpretation of FAR 91.225(e)(2), and how this pertains to the rest of the regulation, see this related answer to a related question: A question about FAR 91.225(e)(2) -- pertains to ADS-B-out requirement and exemption for aircraft with no electrical system .

The truth is that the FAA has simply replaced one poorly-written piece of regulation with another.  In FAR 91.215(b)(3)(ii), the "or below 10,000' MSL" clause was completely irrrelevant.  It is likely that FAR 91.215 was intended to place greater restrictions on flight above 10,000' MSL by aircraft with no electrical systems and no transponders than ended up actually being the case-- for more see related answer FAR 91.215 (b)(3)(ii) -- transponders-- does the "or below 10,000' MSL" clause have any practical significance? .

The new regulation FAR 91.225 suffers from the opposite problem-- not only is it still not clearly stated that the "Below the altitude of the ceiling of a Class B or Class C airspace area designated for an airport" clause of FAR 91.255(e)(2) only applies within the lateral boundaries of said Class B or C airspace, but there is also a possibility of construing the "or [below] 10,000 feet MSL, whichever is lower" clause of 91.255(e)(2) as applying everywhere, not just within the lateral boundaries of Class B or C airspace.  This would have the effect of preventing aircraft without electrical systems and without ADS-B-out from flying ANYWHERE above 10,000' MSL over the 48 contiguous states, unless also below 2500' AGL.  This is clearly not the intent of the regulation, and there has to date been no suggestion that it will be interpreted this way.

This ambiguity could have been avoided by simply writing FAR 91.255(e)(2) to read something to the effect of:  "If within the lateral boundaries of Class B or Class C airspace, then all operations must be conducted below the ceiling of that Class B or Class C airspace or below 10,000 feet MSL, whichever is lower."

One other point which is beyond the scope of the original question, strictly speaking, but may be of interest-- FAR 91.255 will also prohibit any aircraft without ADS-B-out from flying in any Class E airspace at and above 3,000 feet MSL over the Gulf of Mexico from the coastline of the United States out to 12 nautical miles.

Additional links to answers related to FAR 91.215 or FAR 91.225--

FAR 91.215-- may an aircraft with no electrical system and no transponder overfly Wheeler AAF at 9500' MSL?


Can TRACON increase plane speed on approach? I was listening to Norcal yesterday and heard a controller request an airplane increase speed to 210 knots "if able". The plane was previously at ~180 knots. The pilot responded he had flaps 2 and politely declined.

Can TRACON demand a plane increase speed if the controller already assigned a slower speed?
And if yes, what would be the phraseology? "Maintain 210 knots or greater until DUMBA"? (I don't think I've ever heard that when a plane was slower.)

Yes, we can ask pilots to speed up on approach, but it is discouraged and not commonly done. This is because, as an aircraft approaches an airport, the crew will configure it for landing. This includes lowering the flaps, which lowers a number of reference speeds - essentially, making them unable to fly very fast. In order to speed up, they would have to raise the flaps again, which would result in extra workload and bad fuel economy.

The controller you heard knows this. They used the phrase "if able" to emphasize to the pilots, that they only had to comply if it was convenient for them.

ICAO provides the following guidance (form Doc 4444):

  4.6.1.3    Speed adjustments should be limited to those necessary to establish and/or maintain a desired separation minimum or spacing. Instructions involving frequent changes of speed, including alternate speed increases and decreases, should be avoided.

  ...

  4.6.3.5    Arriving aircraft should be permitted to operate in a clean configuration for as long as possible. Below 4 550 m (FL 150), speed reductions for turbojet aircraft to not less than 410 km/h (220 kt) IAS, which will normally be very close to the minimum speed of turbojet aircraft in a clean configuration, may be used.

  4.6.3.6    Only minor speed adjustments not exceeding plus/minus 40 km/h (20 kt) IAS should be used for aircraft on intermediate and final approach.

  4.6.3.7    Speed control should not be applied to aircraft after passing a point 7 km (4 NM) from the threshold on final approach.


What is the history of Concorde motto "May Concorde bring people closer" and medal? I found this image of a medal figuring Concorde and a globe with two contours of continents, before and after Concorde, expressing the new speed of air transport with a motto "May Concorde bring people closer." ("Que Concorde rapproche les hommes").

As I find this material extremely relevant to express the global geography of time-space during the supersonic period, to I would like to know more about it:

Where does this medal and motto come from?
Was it present onboard?
Was it used in other forms of communication?

That medal is a copy of one that was presented to Queen Elizabeth II by M. Ziegler (S.N.I.A.S) during a state visit by the queen to France in 1972, sometime between 15–19 May. The most likely date would have been 18 May, when she visited a SNIAS helicoper factory and was accompanied by Mr. Ziegler.

It was presented to commemorate the intention to build the 'Concorde' for commercial supersonic flying and currently sits in the Royal Collection Trust. "M. Ziegler" is Monsieur Henri Ziegler, who passed away in 1998 at age 91.

"In 1968, he was appointed to run Sud-Aviation, a leading aircraft manufacturer in the south of France, and charged with reorganizing the badly splintered aircraft industry. Sud-Aviation was later absorbed by the larger Societe Nationale Industrielle Aerospatiale, or Snia, which in 1984, as the core of the French aerospace industry, was renamed Aerospatiale S.A."

That is where the "S.N.I.A.S." comes from, "Société nationale industrielle aérospatiale". The Concorde was built by British Aircraft Corporation (BAC) and Sud Aviation. Mr. Ziegler was instrumental to the partnership, and was seen by some, at the time, as a traitor to France for suggesting it.

I think it is fair to say that this motto was coined by Ziegler for the creation of the medal, likely to express his gratitude for England's role in the partnership, but it is not the official motto of the Concorde.

British Airways' slogan for the Concorde was "Arrive before you leave." I honestly don't remember, nor can I find, a French slogan for the Concorde, but the slogan for Air France is "France is in the air."

It was a lot of fun researching the answer for this.


What causes the vortex on the tip of this vertical stabilizer (CFD Simulation) This is a computational fluid dynamics simulation of an airfoil that is oriented to look like a vertical stabilizer. After watching the video, it becomes clear that the tip of the stabilizer causes a counter-clockwise vortex (if looking at the trailing edge). According to this answer, vortices only form on vertical stabilizers if a control surface (the rudder) is deflected. In this simulation, the rudder appears to be completely in-line with the rest of the airfoil.

Furthermore, there also appears to be a slight clockwise rotation on the bottom of the stabilizer, where it would attach to the aircraft's tail. This seems even harder to explain, given the opposite rotation and angular velocity of the top one. I can only assume this is partially caused by the disturbances from what look like hinges farther up. Is this reasonable?

TL;DR: What causes the strong vortex on the tip of this vertical stabilizer?

The subtitle of the video states:

  This video shows a high-fidelity CFD simulation of flow control applied to realistic wing profiles using PHASTA and ParaView Catalyst. Work done by Michel Rasquin from Argonne and Ken Jansen from UC Boulder.

Realistic wing profiles sounds like: we're actually looking at a wing in a vertical orientation - or in a horizontal orientation from a viewpoint that causes some confusion. Looking at a wing at some Angle of Attack that creates lift, and therefore the visualised flow pattern. And then the two vortices are caused by the end effect of finite wing span.


How can an aircraft travel faster than speed of sound without breaking the sound barrier? An excerpt from the BBC article here https://www.bbc.co.uk/news/uk-england-london-51433720

"Despite travelling faster than the speed of sound the plane would not have broken the sonic barrier as it was helped along by fast-moving air."

Could someone please elaborate on how that is possible?

The sound barrier is a function of airspeed and air density. If you have a strong enough tailwind, your ground speed would be above what most lay people would consider the speed of sound without breaking the actual sound barrier.

Aircraft (or any object moving through the air alone) are governed by airspeed. Ground speed is only relevant when traveling on the ground or transitioning to that state (approach to landing).


Why is V1 based only on using brakes? V1 is the speed that an aircraft may abort a takeoff so long as the airspeed is below it, or is committed to flight if it is over that speed.

V1 calculations are based only on the brakes (not sure about spoilers) being available to stop the airplane (regardless of the number of engines), why?

I don't have a proof for the thinking behind the rule making. But my understanding behind it is to prove that the brakes alone can do it in case the reversers are inop. But usually there are exceptions to rules:

dry runways
Dry runway takeoff data does not (with very few exceptions, such as the 737-300/-400/-500 certified to CAA rules) include credit for thrust reverse in the accelerate-stop calculations.
non-dry runways
Historically, Boeing has taken credit for reverse thrust when producing data for non-dry runways. This has been accepted by the FAA. The UKCAA rules, and now the JAA and European certification rules, also allow credit for reverse thrust under these conditions.

For the non-dry rule, Boeing uses half the engines, so for example a 747 would be based on two reversers, and a 737 on one.
Spoilers on the other hand are required so that the landing gears are compressed, if they weren't the tires will slip more. It is also true the heavier the aircraft, the better the brakes work up to their limit.

Quotation and image source: Jet Transport Performance Methods.


What regulatory load cases drive the structural design of aircraft? What I am looking for here is a sort of a list gathering the manoeuvres and/or load cases imposed by regulation that lead to the sizing of a part of an aircraft. For instance, vertical gust load is often the load case that has to be used to size the structural parameters of the horizontal tail (bending moment). What are the other load cases and manoeuvres that usually size the subparts of a commercial aircraft ?

The applicable US regulation is FAR 25.301-563. This obviously goes into a lot of detail about the different aspects that should be taken into account. I will try to condense the cases as much as possible, but many are conditional and relate to various other conditions. Many cases include some subjectivity as to what the FAA will require. Every certification is unique.

While the requirements are somewhat broken down into the general areas of the plane, it's impossible to say exactly which areas of the plane will be critically affected by a certain case. The impact depends on the design and configuration of the specific plane.

There are the upper level requirements:

  (a) The structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.

  (b) The structure must be able to support ultimate loads without failure for at least 3 seconds...

  (e) The airplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to VD/MD, including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope.

Definitions of various important values (which will have structural implications):

VA - Design maneuvering speed
VB - Design speed for max gust intensity
VC - Design cruising speed
VD - Design dive speed
VDD - Design drag device speeds
VF - Desgin flap speeds
VMC - Minimum control speed with critical engine inoperative
VS0 - Stall speed with landing flaps
VS1 - Stall speed, flaps retracted
CN - Normal force coefficient

All strength requirements must be met at all places within the maneuvering envelope, particularly A-I on the following V-n diagram:

Max pitch control deflection at VA (25.331)
Specified pitch control displacement (25.331)
1.5g pullup (25.335)
Discrete gust, vertical and lateral (25.341)

VC: varies linearly

+/- 56.0 ft/sec - sea level
+/- 44.0 ft/sec - 15000 feet
+/- 26.0 ft/sec - 50000 feet

VD: 0.5 times VC values

Continuous gust, vertical and lateral (25.341)
Fuel/oil loads from zero to maximum (25.343)

Addtional factors if there is a "structural reserve" of fuel considered

High lift devices (25.345)

If flaps are to be used for takeoff, approach, or landing, limit load factor of 2.0

+/- 25 ft/sec normal to flight path

Separate conditions, limit load factor of 1.0

Propeller slipstream from max continuous power at VF
Propeller slipstream from takeoff power at >= 1.4 times VF and max weight
Head-on gust of 25 ft/sec

If flaps are to be used en route:

Maneuver to positive limit load factor
Discrete vertical gusts

Maneuver load factor of 1.5 g at max takeoff weight, flaps in landing configuration

Roll maneuvering (25.349)

Airplane load factor 0 and 2/3 of positive maneuvering factor

Steady rolling velocities
Maximum angular acceleration
VA Sudden deflection of aileron to the stop
VC Aileron deflection for same roll rate as VA case
VD Aileron deflection for at least 1/3 of VA case

Unsymmetrical gusts

Yaw maneuvering (25.351), speeds from VMC to VD

Sudden max rudder deflection (conditional definition)
At that max deflection:

Yaw to overswing sideslip angle
Yaw to static equilibrium sideslip angle

Rudder suddenly returned to neutral

Engine torque (25.361)

Limit engine torque at takeoff power and prop speed, 75% limit loads of V-n condition A
Limit engine torque at max continuous power and prop speed, 100% limit loads at V-n condition A
For turboprops, previous conditions with prop control system malfunction
For turbine engines:

Torque from sudden engine stoppage
Torque from max engine acceleration

Engine/APU side loads (25.363)

Limit side load factor

Pressurized compartments (25.365)

All flight loads in addition to pressure differential 0 to max relief valve setting
External pressure distribution in flight
Landing loads combine with max pressure allowed at landing
Pressure including safety factor for high altitude aircraft (>= 45,000 feet)
Sudden release of pressure for critical components

Also affects bulkheads, floors, and partitions

Other emergency depressurizations

Engine failure unsymmetrical loads (25.367)

Failure of critical engine (for turboprops, also failure of drag limiting system)

Speeds from VMC to VD, due to fuel flow interruption
Speeds from VMC to VC, due to mechanical failure

Includes pilot corrective action

Gyroscopic loads (25.371)

Must be considered at max appropriate RPM for certain other cases

Speed control devices (25.373)

Symmetrical and yaw maneuvers and gust conditions, each setting with max speed

If automatic/load limiting features, include corresponding speeds/positions

Control sufrace/system loads (25.391)

Considering other limit load cases

Loads parallel to hinge line
Pilot effort effects
Trim tab effects
Ground gust conditions
Unsymmetrical loads
Auxiliary aero surfaces

Landing load condtions (25.473)

Descent at 10 fps, design landing weight
Descent at 6 fps, design takeoff weight
Both cases should be considered in the following cases:

Level landing
Tail-down landing
One-gear landing
Side loads

Rebound landing
Ground handling
Taxi, takeoff, landing roll
Braked roll
Turning
Tail wheel yaw
Nose wheel yaw/steering
Pivoting
Reversed braking
Towing loads
Unsymmetrical ground loads

Seaplanes (25.521)

Step, bow, stern landings
Unsymmetrical landing
Hull/main float bottom pressures
Auxiliary float loads
Seawing loads

Emergency landing (25.561)

Occupants must have reasonable chance of escaping serious injury in minor crash landing if occupant experiences the following forces:

Upward 3.0g
Forward 9.0g
Sideward 3.0g on airframe, 4.0g on seats
Downward 6.0g
Rearward 1.5g

Large masses should not:

Cause injury to occupants
Penetrate fuel tanks
Block escape

Protection of occupants in dynamic conditions
Ditching strength

Fatigue and damage tolerance (25.571)

Residual strength with certain failed structure:

Limit symmetrical maneuvers
Limit gust conditions
Limit rolling conditions
Limit yaw conditions
Pressure differentials
Limit ground loads

Safe-life evaluation, if failure conditions are not practical
Sonic fatigue
Safe flight with damage from:

4 pound bird at VC at sea level and 0.85VC at 8,000 feet
Uncontained fan blade impact
Uncontained engine failure
Uncontained high-energy rotating machinery failure


Load testing a model glider wing I wish to load test a model glider wing.

The standard static load test methodology using sandbags on the upturned wing would be difficult as the chord of the wing is small - it will be difficult to securely attach the loads required.

Are there any other methods commonly in use to do this kind of test that might be easier?

You can still use some small bags with flour or other soft material inside.

For the wing load test, the bending moment is more critical than the torsion. Therefore, the flour-bags could work fine.

If you want to add torsion and a more precise loading, you can build some sectional fixtures that transfer point loads to the wing.


Did the Douglas SBD Dauntless have a second set of controls? In the book Never Call me a Hero, the author mentions that there were controls in the rear seat of the Douglas SBD Dauntless. However, I don't recall ever hearing this before, though I know the TBD had controls in the second seat. Did the SBD have controls in the rear seat?

The Dauntless did have controls for the rear gunner- a pair of permanently attached rudder pedals and a control stick that was stowed away on the port side until needed. See:

https://www.tapatalk.com/groups/warships1discussionboards/sbd-dauntless-flyable-from-rear-cockpit-t2435.html

Juan Jimenez is correct that it would be very difficult to operate, what with the armour plate for the pilot and the radios right in front of the gunner. See here for a story of flying a veteran Dauntless where the pilot gets a brief familiarisation flight from the gunner's seat (it's mentioned in the fourth paragraph after "SBD Glider: Engine Out"):

https://www.aopa.org/news-and-media/all-news/2011/january/pilot/flying-a-national-treasure-the-sbd-dauntless

You can see the gunner's control view at the Smithsonian's walk-around view of the Dauntless, it's images four and seven in the collection (six shows the pilot's view and eight the gunner's real place of business):

https://airandspace.si.edu/collection-objects/douglas-sbd-6-dauntless

Photo seven


How does the forward speed of an auto-gyro relate to the rotation speed of the rotor during auto-rotation? In my other question I was looking for a relation between the forward speed and the thrust (lift) of an auto-gyro.

I have come to the conclusion that for an auto-gyro in auto-rotation the forward speed will rotate the rotor and with that regulate its rpm. The faster an auto-gyro flies the faster the rotor rotates resulting in a higher thrust.

What I hope to accomplish is forward braking where the auto-gyro flies faster than needed to provide sufficient thrust, resulting in a higher rotation speed which will then result in more thrust and lift the auto-gyro up. This additional thrust isn't needed as the same level is required. By braking the rotor, using a generator, the rotor speed is reduced till the point it is still able to maintain level using auto-rotation. For this theory I need to understand the relation between the forward speed and the rotation speed.

But it is still not completely clear how this forward speed relates to the rotation speed. I hope someone can explain it to me or point me into the right direction.

I might be completely wrong and came to the wrong or inconclusive conclusions if so don't hesitate to tell me.

Thanks in advance!

An autogyro flies at constant rotor RPMs. In vertical autorotation, that is a particular case, the rotor revs may be 10-15% less, but in straight and level flight, the rotor revs are higher, and always the same, independently of the speed of the gyro. It cant be otherwise, since, if the rotor revs went higher with the speed, the rotor thrust would also rise, and the gyro will gain altitude, no longer flying straight and level...

It's also true that if, flying s/l, the pilot opens the throttle, he also pushes the stick slightly forward in order to reduce the AoA of the disk, and thus keep flying s/l...


How are modern fighters cleaned? I found this photo of a fighter jet being cleaned, how does this process work?

(usatoday.com)

Yes, they use a chemical to clean aircraft. Back when I was in the Air Force, we used PD-680 type II.

I'm pretty sure they don't use that any more though. PD-680 was discontinued (at least for most applications) around the turn of the century. Most applications that used to use PD-680 switched to PRF-680. I can't guarantee it, but I'd guess that's what they use to clean aircraft.


In a helicopter does forward flight produce a rolling moment? In forward flight the advancing blade will produce more lift than the retreating blade. It seems this would produce a rolling moment. If so, does the pilot have to move the cyclic to compensate for this? If it doesn't, why not?

Yes cyclic is used to compensate.

The advancing blade is allowed to "flap" up to compensate for its extra lift, and the retreating blade flaps down to compensate for its decreasing lift.  On an articulated rotor, each blade has a flapping hinge for this purpose.  On a two blade teetering rotor, the blades being on the same axis are flapping up and down in harmony and the teetering pivot at the center takes care of it.

If the cyclic was perfectly centered (swash plate horizontal) the rotor disc would be flying a long with a slight tilt away from the advancing side as the machine flies straight.  But this would impart a lateral moment that would make the machine turn to the left as it flies forward.  So to compensate the pilot is initially applying a small amount of compensating right cyclic without thinking about it (on a machine where the right side advances), so effectively the advancing blade gets a slight reduction in pitch and the retreating blade gets a slight increase in pitch, becoming more and more the faster you go.

As you get faster and faster, either the retreating blade runs out of pitch authority or wants to stall as its relative speed gets too low, or the advancing blade's tip becomes trans-sonic. One or the other.  This is what limits helicopter speeds.


Do the wheels of taxi-capable helicopters have power? There are many big helicopters which seem to taxi just as good as any fixed wing aircraft, for example CH-47.

Do their landing gear wheels have power or they also taxi using thrust like their fixed wing counterparts?

If they do have power, are the taxi controls different from the standard controls?

The wheels do not have power - there is no propulsion system linked to the wheels, they just spin freely.

Taxiing is achieved by adding just enough collective so that the helicopter starts moving, but not enough to lift it off the ground. A little forward cyclic also help.

To slow down, the main wheels have brakes, like fixed-wing aircraft.


Can two Airlines have the same ICAO code but different IATA codes or vice versa? We are trying to import all the Airlines into a Table.
Plan is to Update existing Airline Data if not present then create it.
For this we needed an unique way to identify the Airlines.
Is ICAO+IATA an unique identifier or is it possible to have Airline without IATA or ICAO code?

IATA codes are not unique. There simply aren't enough of them and they are expensive to maintain. Additionally the concept of an airline is a little fluid.

The simplest issue to deal with are controlled duplicates, where 2 airlines that are geographically seperated share an IATA.

Another key issue is where an airline may have one commercial brand but several corporate entities each with there own IATA for cabotage purposes. LATAM group is a key example of this, but several european airlines (easyjet and ryanair) are also going down this route to be prepared for brexit.

Many small regional airlines in places like Alaska,the Australian outback and the Caribbean will simply not be able to afford an IATA code, or may make up there own 2 letter code that is not an IATA code.

Finally you have franchise operators like flexflight, who operate a single route themselves to maintain membership, and then provide ticketing services to other small airlines, using there code to interact with the GDSes. In this case, one IATA code can represent many potentially geographically overlapping airlines, but they are usually separated by something, often flight number.

If you wish to track all airlines in general, you are going to have to come up with your own id.


Stress on wing/fuselage-joint of aerial water tanker aircraft - and possible mitigating measures Following the unfortunate crash of a C-130 aerial water tanker aircraft in Australia, a friend brought to my attention a 2002 C-130 tanker crash in the US. As can also be seen in this video, the wings fold upward and separate from the fuselage, which then rolls over and crashes.

According to the referenced Wikipedia-article, citing the NTSB (1, 2), the wing/fuselage-joint (likely) failed due to fatigue:

  The NTSB investigated the crash and determined that the accident was caused by a structural failure that occurred at the wing-to-fuselage attach point, with the right wing failing just before the left. The investigation disclosed "evidence of fatigue cracks in the right wing's lower surface skin panel, with origins beneath the forward doubler. The origin points were determined to be in rivet holes which join the external doubler and the internal stringers to the lower skin panel. These cracks, which grew together to about a 12-inch (30 cm) length, were found to have propagated past the area where they would have been covered by the doubler and into the stringers beneath the doubler and across the lap joint between the middle skin panel and the forward skin panel."

This and another crash in the same year subsequently led to a fleet grounding, partially following concerns over training and maintenance.

Besides the incident referred to above, this got me thinking. The substantial and rapid change in weight when the tanker performs the drop must result in (temporary) large forces/moments on the wing/fuselage-joint. According to the summarised NTSB-report, the aircraft in the video carried some 3000 gallons of retardant (some 7000kgs at 9lbs/gallon), but does not specify volume flow. A more recent Modular Airbore FireFighting System that fits into a C-130 can disperse something under 3000 gallons in 5 seconds, yielding 600 gallons/sec - equivalent to 5400lbs or 2450kg/sec. Various comments on the YouTube-video (such as this and this) also point in that direction -- but unfortunately don't include sources. As such, I'm wondering:

to what extent these changes in weight indeed introduce forces/moments that (might) go beyond the design specifications of the aircraft?
what (technical or operational) measures are/were normally taken to prevent  such problems (structural reinforcement during the conversion, limiting 'offloading speed', no dropping during climbing/manoeuvering, etc.)?

The problem isn't the relief of stress from dumping of loads.  If it was, you would have the same failures happening because of the act of landing, where the wing goes from being heavily loaded to unloaded in about the same amount of time.  And it's once per flight in any case and is something that all airplanes have to cope with.

Fatigue is all about the number of load cycles, the beefiness of the structure designed to handle it, and identifying the "weak links in the chain" stucturally.  The problem for tankers is they spend a large amount of time being subjected to very sharp  and relatively numerous gust loads during drop operations.  Working a fire, they are flying through extreme thermals and mechanical turbulence; it's a pretty rough ride.

So the airframes get "beaten up" you might say.  Effectively it's like taking a car that could handle the odd pot hole with no problem and driving it exclusively on a road full of potholes all the time.  Sooner or later something will break on the suspension, if you didn't anticipate the effects beforehand. And in the case of the C-130s that's about the long and short of it - they didn't anticipate the effects beforehand and the airplanes were already old.

So what happens is, the more extreme operating environment means the machine is living much closer to the margins of the original structural design analysis. This raises the risk that flaws in the original structural fatigue analysis will float to the surface, when they would normally stay dormant. Structural inspection programs that were working fine when the airplanes were being used in their normal role are suddenly inadequately frequent and/or comprehensive, and cracks form that were never a problem before, aren't being inspected for, and... ooops.

Unless you want to stop using them to fight fires, the answer is pretty simple.  These are the options:

Inspect more frequently, and or more invasively using NDT procedures.
Add more meat, or redundant meat (damage tolerance) to the structure.
Both (the probable result).


On IMA systems e.g. A380 how many people are required to preform regular maintenance checks on this system? I am aware that there is usually 1 or 2 engineers preforming maintenance checks for a particular avionic system on an aircraft at a given moment. For IMA system, is there often more engineers working to get the maintenance checks completed. Are the maintenance check procedures/tests on IMA systems more complex that would require more people.

Not any more people than any other system. But your question raises a few key points for consideration.
First of all, maintainers don't actually go anywhere near the aircraft if it is not a requirement. I don't agree that there are usually 1 or 2 engineers on an aircraft at any given moment. If it aint broke, don't fix it. Over-maintenance is a cause of systems breaking more often due to factors such as human error, more cycles that could reduce life, etc. So unless there is a reason to be there, they won't be pushing buttons just to see if something is still working. In fact, your foreman/chief would probably ask you why you were touching that aircraft. Valid reasons to be working on the aircraft are most often that the maintenance is called for in the servicing schedule (scheduled maintenance), or it has broken (unscheduled maintenance). There are other exceptions to this including modifications and special technical instructions. These will have slightly different names depending on the local authority under which the aircraft operates.
Second point is that if you have been granted authority, have the current licence and the company policy allows for it, you may be able to perform maintenance by yourself. However, airliners usually (at least in Australia) have an AME (aircraft maintenance engineer) do the work and sign the first level of certification, followed by a LAME (Licenced Aircraft Maintenance Engineer) who certifies that the work was completed in accordance with the approved procedure and issues a certificate of release. This allows the aircraft to go flying again. In general aviation, it is more common for the LAME to do the work by themselves. In this case, they are kind of supervising themselves.
Third point: The complexity of the aircraft's system does not determine the number of "levels" of certification. If anything, it is how safety critical the system is to airworthiness that determines how many people must sign for the job. In some cases, an IMI (Independent Maintenance Inspection) or QA (Quality Assurance Inspection) is called for, increasing the levels to 3. Systems such as landing gear are pretty important so they may have 3 levels, but this isn't always the case. The levels are 1-AME, 2-LAME, 3-Independant. I don't think I've heard of Integrated Modular Avionics requiring 3 levels. As a training manager, I'd just make sure my people have adequate training and are confident. If they aren't confident, part of getting your LAME licence is knowing when to put your hand up and say you can't do something. Much safer to do that than guess what you are doing is correct. Hundreds of lives are at stake. You need to have the knowledge and maturity to be a good engineer.
Another interesting point worth bringing up: The Royal Australian Navy had a Sea King helicopter crash leading to loss of lives during Operation Sumatra Assist. This was for the 2004 tsunami. The cause was the lack of a split pin that should have been installed on the main rotor hub assembly. Long story short, the Navy increased the levels of inspection of this job (and others similar to it) from 3 to 4. Data showed that after this change was implemented, more mistakes happened, which is a bit counter-intuitive. You would think that by having 4 people look at something that it would be safer. Wrong. Data showed that people became more complacent because they thought that someone else would have their back if they were having a bad day. The Navy reverted back to 3 levels, and incidents reduced again.
IMA systems (I assume you mean Integrated Modular Avionics) aren't any more complex than aircraft without it. For anyone who is wondering what IMA is, it is basically a concept of merging what used to me a bunch of separate computer systems into a reduced number of boxes. That way the systems can share computing power. On the 787 these are called CPIOMs (Central Processing Input Output Modules). They've managed to reduce the physical footprint and weight of the entire avionics package than pre IMA concepts. There's actually a good argument to say that it is now simpler, because there are less boxes to maintain than before. System testing (BITE) is still done via an interface in the cockpit and is as easy as before.


What is the reasoning behind using depleted uranium as counterweights in the 747? I recently found out that some of the earlier models of 747 had between 600 and 1100lbs of depleted uranium as counterweights. These counterweights were found in 747s manufactured between 1968 and 1981.

(Source)

Given that the study linked above and the well-known radiation exposure for workers (and passengers) even prior to the 1968 introduction of these counterweights, why did Boeing (and McDonnell Douglas) decide to use such a hazardous (and potentially dangerous from a military perspective) material?

I understand that the weight/density allows for a smaller package, but it seems that the tungsten weights were integrated after 1981 without any space issues in existing aircraft.  Was depleted uranium in such high supply that it was also more economical?

Eventually, the counterweights were replaced with these tungsten replacements:

(Source)

Boeing

Boeing would have used DU because it had the right combination of physical characteristics and cost. Their tests showed that the radiation exposure for workers was low (2.6% of the statutory "safe" level). In most cases the exposure was so low as to be not detectable.

Radiation

Passengers on aircraft are exposed to cosmic radiation at much higher levels than people on the ground.

On the ground the average American is exposed to 620 mrem/year from all sources.

Those flying from Washington DC to Los Angeles would be exposed to close to 2 mrem from cosmic radiation. This is not an issue for passengers but is something that airline crew and other very-frequent flyers are aware of.

Exposure to aircraft occupants from cosmic radiation is 600 times higher than from DU counterweights in the early 747 aircraft.

Depleted Uranium

DU is Uranium from which the most radioactive parts have been removed. It is less radioactive than naturally occurring Uranium.

Heavy metals like Lead, Tungsten and Uranium are toxic if ingested. This toxicity is not due to radioactivity. Aircraft crew and passengers are not exposed to these metals in any way that could cause them to be ingested.

Boeing Tests

Boeing carried out safety tests on the material.

Workers

  Boeing
  has
  conducted
  two
  dosimetric
  studies
  of
  exposures
  to
  workers.
  For
  both
  studies,
  whole
  body
  exposures
  were
  measured
  with
  film
  badges
  which
  were
  provided
  and
  processed
  by
  Landauer,
  Inc.
  These
  badges
  have
  a minimum
  detectable
  exposure
  of
  10
  millirem
  per
  issue
  period
  (monthly).
  In
  the
  second
  study,
  extremity
  exposures
  were
  measured
  with
  finger
  rings,
  also
  provided
  by
  Landauer,
  having
  a
  minimum
  detectable
  exposure
  of
  30
  millirem
  per
  issue
  period
  (also
  monthly).
  The
  periods
  of
  the
  study
  were
  December
  1968
  to
  February
  1970
  for
  the
  first
  study
  and
  September
  1977
  to
  April
  1978
  for
  the
  second.

  Both
  of
  these
  studies
  showed
  all
  worker
  whole
  body
  exposures
  were
  less
  than
  2.6%
  of
  the
  exposure
  limits
  for
  occupationally
  exposed
  employees
  (5000
  millirem
  per
  year)
  and
  less
  than
  26%
  of
  the
  limits
  for
  members
  of
  the
  general
  public
  in
  effect
  at
  that
  time
  (500
  millirem
  per
  year).

Crew and Passengers

  Based
  on
  the
  data
  from
  National
  Lead,
  reported
  in
  the
  section
  on
  exposures
  to
  the
  general
  public,
  dose
  rates
  to
  flight
  crew
  will
  be
  less
  than
  0.8
  microrem
  per
  hour.
  During
  a
  2000
  hour
  working
  year,
  this
  results
  in
  a
  maximum
  potential
  exposure
  of
  1.6
  millirem,
  less
  than
  2%
  of
  the
  100
  millirem
  per
  year
  limit
  for
  members
  of
  the
  general
  public.
  This
  is
  only
  1/600th
  of
  the
  500
  microrem
  per
  hour
  increase
  in
  dose
  rate
  from
  cosmic
  radiation
  flight
  crew
  experience
  at
  39,000
  feet.

Boeing Use of Depleted Uranium Counterweights in Aircraft. 1984. Retrieved from NRC

References

Space Radiation, Should Frequent Flyers Worry? 2013. BBC
Tungsten Poisoning. 2014. Retrieved from Patient.info


How are backlit aircraft panels (for simulators) made? I'm wondering by which method this sort of aircraft panel that enabling backlighting are made? What is the process? I've gathered they are acrylic but how is the gray paint selectively removed?

pic http://www.simparts.de/WebRoot/Store/Shops/15465782/4B72/6E37/0BF8/50E9/641B/C0A8/28BA/29AA/A320-OHP-009-5.jpg

This video says the following:

  Here's my Tonson TS-3040C CNC making an EFIS panel for my flight
  simulator. For the lettering I am using a 0.4mm End Mill and for
  cutting out I'm using a 1mm End Mill. RPM is 10000 and movement is
  300mm / Min. The material is Cast Acrylic which I have given one coat
  of black paint and two coats of grey paint.

So that's at least one method, but there are probably more...


What precautions are taken to avoid bird strikes? According to this Wikipedia page on bird strike,

  The International Civil Aviation Organization (ICAO) reported 65,139
  bird strikes for 2011-14, and the Federal Aviation Authority counted
  177,269 wildlife strike reports on civil aircraft between 1990 and
  2015, growing 38% in 7 years from 2009 to 2015. Birds accounted for
  97%.

What precautions are taken so that birds are not killed by aircraft?
Is the pilot held responsible or does the pilot pay some penalty in case innocent animals are killed?
Are there any planes which have been designed specifically focusing on bird safety (or have there been any efforts in designing planes focused on bird safety)?
Where can I get reports on the number of bird strikes per year on an international level?

Note: This question does not concern the damages the airline suffers from bird strikes.

Update In Addition To Other Answers

Why jet engines are NEVER protected in the front?!

Humans are more important than animals by "most" societal standards. Therefore, bird safety is secondary to human safety (pilots and passengers). However, there are certain areas (in US) that are designated as protected areas for certain types of birds, such as in California where I have seen protected areas for the California Condor which requires pilots to fly at a certain altitude above that area. On VFR Sectional Charts the special conservation areas are depicted as indicated here:

These conservation area restrictions are mostly to avoid disturbing the birds, but I assume it also reduces the risk of a bird strike by requiring the aircraft to fly higher over these areas.

One of the most important precautions is for pilots to remain vigilant and see and avoid, but that is not always possible when a flock of thousands of birds takes off near an airport when airplanes are landing and taking off.

Some airports use various methods (see links provided) to scare flocks of birds away such as using hawks, noise, and other methods.

I don't know of any aircraft that are designed for the safety of birds. I believe this is a very low priority in aircraft design.

Try checking the Audubon Society website, they have a lot of information on risk to birds and statistics on such risks.


Can 737NG GPWS and other warning voice lines/sounds be played in headphones instead of cockpit speaker? I watched a couple of cockpit videos on the 737NG landing, and I noticed GPWS in some of the videos does not announce 500ft, 400ft, 300ft, etc. Does this mean the pilots have routed these sounds into the headphones, or they have been disabled on the particular operator’s aircraft?

No, the pilots cannot disable audio warnings or switch them over to headphones only. They are always send to the speakers and all headphones. The pilots cannot even adjust the volume for these warnings. This is to make sure that these warnings can always be heard regardless of any intentional or accidental misconfiguration of the audio control panel.

  Audio warnings for altitude alert, ground proximity warning, collision avoidance,
  and windshear are also heard through the speakers and headsets at preset volumes.
  They cannot be controlled or turned off by the crew.

(Boeing 737 NG FCOMv2 5.20.1 - Communications - System Description)

The typical altitude callouts from the GPWS are:

  Approach Callouts

  Radio Altitude Callouts

  [Option - Typical]

  The GPWS provides the following altitude callouts during approach:

  2,500 feet – TWENTY FIVE HUNDRED
  1,000 feet – ONE THOUSAND
  500 feet – FIVE HUNDRED
  100 feet – ONE HUNDRED
  50 feet – FIFTY
  40 feet – FORTY
  30 feet – THIRTY
  20 feet – TWENTY
  10 feet – TEN.

  Note: Callouts at 1000 feet and 500 feet are based on barometric altitude above
  the landing field elevation; callouts at 2,500 feet, and below 500 feet are
  based on radio altitude.

  DH/MDA Callouts

  The GPWS provides height callouts based on the altitude set by the Captain’s Minimums selector.

  Callouts are based on radio altitude when the MINS selector is set to RADIO. Callouts are based on barometric altitude when the MINS selector is set to BARO:

  DH/MDA plus 100 feet – PLUS HUNDRED
  at DH/MDA – MINIMUMS

(Boeing 737 NG FCOMv2 15.20.17 - Warning Systems - System Description)

All of these callouts are airline options and can be disabled by individual operators. There are also callouts at other altitudes, which are not listed here. See also Does the Boeing 737 NG have "5" radar altitude callout?.


Why do delta wings generate less lift at a particular angle of attack compared to straight wings? I simulated a delta wing and a straight wing in xflr5 with the same profile, wing area and airspeed (very low airspeed. around 10 m/s. it's for a model aircraft). The straight wing was operating at nearly the lift coefficient of the 2d airfoil at a particular angle of attack, but the delta wing was operating at a lift coefficient far below the 2d airfoil coefficient. Why is this so? The wing has a medium sweep of 25 degrees. Will the lift improve with a higher sweep angle?

No, a higher sweep angle will reduce lift even more.

All aerodynamic forces scale with the dynamic pressure, the product of air density and the square of airspeed, divided by half. If a wing doesn't produce lift, move to denser air or speed up.

However, drag will also scale with dynamic pressure, and there is more which influences lift. One is the angle of attack and the second is the suitability of the chosen wing geometry to produce lift at a given speed and angle of attack. Straight wings score best in this respect, at least at subsonic speed.

The sweep angle of the quarter-chord line gives you a first approximation of lift loss. Lift is proportional to the cosine of that angle. Just drive sweep to its extreme of 90°, and the inability of such a wing to create lift should become clear.

Lift is produced by the downward deflection of the air streaming around a wing. When the wing meets this flow straight on, it will have the biggest effect on the air and create the most lift. Any sweep, regardless of forward or backward, will reduce the wing's effect.

The next effect is the influence of the wingtips. If your wing extends sideways a lot more than it extends in flow direction, the influence of the tips is relatively lower. The ratio between wing span and chord is called the aspect ratio, and it figures prominently in the formula for the lift curve slope.


What would happen if the blades of a variable pitch prop did not all have the same pitch? What would happen if one or more of the blades (but not all) on a variable pitch prop would not tilt at the same angle as the other ones?

Could this situation happen or is the technology so advanced or stable that it would be basically impossible?

Should it happen, besides vibration (most probably due to imbalance factors,) what other effect could there be? How would pilots solve this emergency (if it can be called so?)

Never say never, but that's not a prop malfunction that I've heard of or that was ever addressed in the turboprops that I've flown.  We had various prop malfunctions, but the ways that things could fail that we considered would pretty much leave all blades in the same place.  So without claiming that it could not, ever, ever, no way no how, not in all possibility ever happen, I'll say that it's not very likely.

If you did get something like that, the vibration would probably be bad enough that you'd shut the engine down, and at that point the interesting question would be, do all the blades feather for you?  (The answer to that probably depends on what type of prop you have and what specific failure mode led to the condition in the first place, and now we're into too much speculation for any answer to be very enlightening.)


What are the advantages of dual counter-rotating rotors vs coaxial rotors? It just seems to me, as though, dual counter-rotating rotors are a much better solution (Chinook) than coaxial and that the center of gravity is offset to allow useful loads to be spread out. Even though that the airframe needs to be longer, does this give more stability on the whole airframe and if so, would engine failure allow for proper auto rotation? I am thinking of small two-four person model for this question.

Counter-rotate Picture above from this answer, which compares the Chinook with the V-22 Osprey, another counter rotating twin rotor.

For a given weight, the larger the rotor the less power is required for generating the necessary thrust. And here is where a single rotors' advantage is, one single rotor requires less power than two rotors with half size each. In the hover, the single rotor throws away about 10-20% of the engine power to drive the tail rotor, but in cruise flight much of that torque compensation is generated by the vertical tail.

Yes the payload location range of the Chinook config is without equal. In case of one engine failure, the remaining engine can drive both rotors since they are coupled by a shaft. Autorotation is not an issue either - from this thread:

  Chinook is extremely docile in handling, yet extremely maneuverable when light. Autorotation is of the thistledown type, rotor inertia is massive and the beggar wont come down fast even if you want it too. Engine off landing is merely a continuous flare to a running landing on the rear wheels which requires quite a lot of runway. I dont know if the military use other techniques.

So in short, the Chinook config has pros and cons. The two rotors' wakes intermingle with a reduced power efficiency in forward flight. From same thread as above:

  Due to the dynamics the Chinook develops maximum translational lift in sideways flight to the left, hence the antics of the logging machines plugung sideways on the end of theit long cables - gives them greatly increased lift capacity.

So yes the twin counter rotating rotor does not spill tail rotor power during the hover and has a large payload range. It has a longer but narrower footprint than a similar MTOW single rotor. But it is not universally more power frugal than the single rotor config.

Co-axial. Picture above from this question.

The co-axial setup main drawback is its complexity. Counter rotating shafts, one in the other. One rotor quite high over the other one to allow relative flapping. Super complicated control actuation.

Power requirements: the coaxial configuration has power requirements according to this answer.

All in all, the counter rotating setup is simpler than the coaxial setup, which does have the advantage of a very small footprint. But the single rotor set-up has proven to be the configuration with the best design balance.


Is there a minimum number of people needed for an aircraft pushback? This seems a bit wasteful. Is there a legal requirement for a minimum number of people and if so, on what basis does this vary? (airline, airport, aircraft type, etc)

Depends on the airline.  At minimum, you need one driving the tractor, and in unconfined areas with nothing to run into, this is sufficient.  In confined areas operators almost always require wing walkers to clear the tips, and someone at the tail if backing into a confined space or to make sure the pushback stays within any boundaries.

So the answer is, anywhere from 1 to 4, depending on the situation and the airline's internal policy.  Insurance requirements will have an influence on this in a lot of cases because the results of screw-ups are usually just insurance damage claims with nobody getting hurt.


Why would Air France fly to Algeria when it had a civil war? I was reading about the Polaris Award that was awarded to the pilots of Air France Flight 8969:

  Algeria was in a state of civil war at the time of the hijacking. Aircraft flying to Algiers faced the possibility of missile attacks. As a result, Air France's flights to Algiers had crews entirely made of people who volunteered for the route. Air France had asked government officials if it absolutely had to continue flying to Algeria; as of the time of the hijacking, but they never gave Air France a reply. Bernard Delhemme was the captain of the flight.[3] Jean-Paul Borderie was the copilot, and Alain Bossuat was the flight engineer.[1] The Airbus A300B2-1C, tail number F-GBEC, had first flown on 28 February 1980.[4]

Please see the question in the title.
Why would rational air crew volunteer to fly to a state with civil war, like Algeria? Did they have optimism bias? How did they underestimate the risk?

A civil war yes, but it still had a government and this was supported by France.
A flag carrier like Air France does not really have all that much leeway in terms of what they can do and it was probably in the nation's interest to show support to the legal government. Since they had asked, the answer was probably yes they had to fly(but no politician worth his salt would be on record as stating that! Without an airlink to the outside world a lot of things would come to a halt. I would imagine a lot of people who were trying to assist in finding a conclusion (mediators) would be users of the flight. There would also be a steady stream of people leaving the country and AF may be one of the few options available.
Most airlines treat flights into danger-area differently and where unions are strong having volunteers to operate the flight may be an acceptable solution to both the airline and the union.


What is the data size of Hilton Software's WingX Pro 7 app and all of its databases? What is the data size of the WingX Pro 7 installation files and databases including the app itself and the VFR Sectional, IFR Low Enroute, IFR High Enroute, and the full "All Charts(IFR)" databases?

I've noticed that the databases for different aviation apps have different sizes. What is the total storage for WingX Pro7 if you select the option to download all of the following databases:

List item
VFR Sections
IFR Low Enroute
IFR High Enroute
All Charts (IFR) - USA

My understanding is that WingX Pro7 uses higher compression than foreflight, leading to smaller database sizes, but I have been unable to find this information on their website, nor have their customer service responded to my email question.

UPDATE: I got my hands on a 16GB iPad Air and installed the latest version (July 2017) of WingX Pro 7 and downloaded all of the databases for the US including VFR and all IFR charts and plates for the Contiguous USA, plus terrain information. The only database I excluded was the High altitude IFR enroute charts. The total storage used by WingX Pro7 on the iPad is reported as 4.62GB.

2nd UPDATE: I finally got a response from HiltonSoftware on this question. They responded as follows:

Support  07/05/2017 at 2:25 PM
To: Devil07
Message body
WingX, fully loaded required 8GB
iOS may require at least 2GB free memory when all App are loaded
Space also has to be allowed for iOS

Support Team

The app itself only takes up about 15-20MB, according to the Google Play or iTunes app store, where as ForeFlight takes up 226MB (almost a quarter of a gigabyte) and Avare is less than 2MB and free. After downloading a good handful of charts, Avare is about 250MB. If you are very tight on space and running Android, you could try Avare. Otherwise if you need WingX Pro 7 or are on Apple, 5-6GB is what you should have free if you plan to use all the charts. If you're buying a new tablet, try to go for a 32GB model to be safe and have plenty of room for other apps additionally, otherwise 16GB should also work.


Why don't modern heavy bombers have gun turrets? I know that for a period of time the B-52 had a tailgunner, but nowadays many heavy bombers lack any turret defense.Why?

There are really two reasons. First, fighter tactics have evolved over time, and second, the role of the B-52 has evolved over time.

Fighter tactics

Fighters used to depend (heavily) upon getting fairly close to an enemy, and shooting it down with a machine gun. Over the years, guns have become less and less dominant, and instead fighters tend to use missiles from outside a machine gun's effective range.

Role Change

The B-52 was designed from the beginning to carry nuclear weapons. Nonetheless, it's never been used to actually bomb an enemy with a nuclear weapon.

Early in its life, the B-52 was used to drop conventional bombs on enemies though. During the Vietnam conflict (in particular) B-52s were used to drop thousands of conventional 500 pound bombs on various targets.

The Linebacker II campaign was, perhaps, especially notable. During Linebacker II, 15 B-52s were lost--all of them to surface to air missiles (SAMs) not enemy fighters.

Linebacker II did see some limited success from tail guns as well: two B-52s shot down two MiGs. No B-52 (nor any other aircraft of similar size) has shot down an enemy fighter since.

Now, B-52s act as an aerial platform for launching stand-off weapons. Its normal operation is to fly in low (where it's difficult to detect on radar), and launch from a substantial distance from the target. Since it can be a substantial distance from the target, an enemy would need to defend a much larger perimeter (e.g., a large number of fighters basically blanketing a large area) to stand a decent chance of intercepting it.

Even if we assume an enemy fighter were to intercept the B-52, we get back to the situation with fighter tactics outlined above--they'd be unlikely to get within range of its tail gun anyway.


How do bearings in a jet engine keep water out when it's been cleaned? When a jet engine is running, it's generally immune to any liquid in the incoming air as the liquid would have been evaporated in the compressor. However, when the engine is being cleaned for carbon deposits, cleaning solution mist is injected is fed to the inlet, and the water would have the chance of entering the compressor bleed. Is there any special design in the bleed path to keep the water out, or the water simply couldn't get past the labyrinthine seals around the bearings?

To keep dirt out, bearings with exposed faces have rotary seals on them. Those seals will also keep water out, as long as the bearing is not rotating when wetted and not completely immersed. If immersed for days or weeks, even well-sealed bearings will eventually allow water in.


How was the 'bump' removed from the Airbus P2F? (Sources: airliners.net, YouTube)

How was the 'bump' removed from [the design of] the Airbus P2F?

The A330 Freighter based on the A330-200 features a bump / blister / bulge that raises the nose landing gear to level the cargo deck for easier loading of cargo. A design issue with the passenger A330 is the slanted floor, but it was not a big deal since cargo with feet manages.

  The A330/A340's distinctive nose-down pitch - which is greater than intended in the original design and only emerged when production of the aircraft began in the early 1990s - was a problem that Airbus had addressed on the first version of the A350.

The passenger-to-freighter (P2F) conversion program for both the -200 and -300 lacks this bump. What design change was incorporated into the conversion that fixed the floor level?

The latest information I could find is from October 2009 (8 years ago):

Airbus abandoned the 'bump' modification in favor of a special ground jack to level the deck during loading/unloading.

  Airbus is proposing a piece of external ground-support equipment which would physically lift the aircraft, producing the same effect as the lowered gear in new-build freighters (flightglobal.com).

The A330 P2F features a fully-powered Cargo Loading System (CLS) that is perfectly capable of loading/unloading even the heaviest pallets even with the aircraft "on the tilt". Therefore, no 'tilt remedy' is required.

Flight Global, 15 February, 2012:

  The P2F version will retain the passenger aircraft's geometry and incorporate a powered cargo loading system to enable pallets to be moved "uphill" on the main cargo deck.


How to calculate the forces and moments required to aerodynamically actuate slats? I am trying to make a wing where slats automatically deploy once it hits a higher angle of attack, let's say right before stall.
I need to roughly estimate the forces and moments acting on the undeployed slat.

Here is a sample airfoil with the pressure distribution superimposed:

Source: own work

I plan to operate at a Reynolds number of 100,000 at stall speed, and the preliminary dimensions for the slat position are 10% for the width, 1.5% for the depth and 2% for the gap, all in chord percentage. Refere to the image below for the definition of the dimensions:

Source: Effect of Slot Span on Wing Performance, J.Granizo

I think the next step is to use the following equations for the slat alone:

Source: Fundamentals of Aerodynamics 6th edition

However, I only know the pressure and friction coefficients on the top surface (marked in red line). Do I assume that the entire bottom surface (marked in blue line) is at atmospheric/freestream pressure?

A point in the right direction would be appreciated.

Very interesting challenge and very interesting design approach.

I've read this had been done few times with TE flaps but not with LE slats. If i were you I would take the following approach to figure a way out.

In Javafoil you can model open slat multi element air-foil. The results are not very accurate at your Reynolds number if there is a huge separation bubble though. xfoil does not support this feature at the moment but if you are thinking of using TE Flap your best bet is to use xfoil). Calculate several slat gaps by geometrically altering the input file for a list of angles of attack.
You can easily export the local CPx data to a text file in Javafoil. Then use these data to find out how much of aerodynamic force the slat generates by integrating the local surface slope and CPx values for each slat-gap, alpha pair.
Then this data-set could be interpolated to find out what would be the slat-gap distance for given angle of attack for a given spring constant and substantially to find L/D for each angle of attack.

Hope this helps.
ABCD

PS: One example of automatic TE flap deflection is discussed here. I think LE slat would be much more effective because $Cp_\alpha$ near the LE is much higher than that of the TE


What does 'given rate' mean? I was watching this video:

At around 1:28 I heard the controller clear a plane to descend to a level "given rate". I was searching online but couldn't seem to find what they meant with the term 'given rate'.

Could it be the "level change" or "open descend" button in the aircraft?

The ARTCC Controller is telling a flight crew currently descending to descend to FL150 at their current rate of descent e.g. 2000 fpm, etc. either to expedite their current descent or the controller is satisfied with the current rate of descent to traffic separation and flow into the terminal area where the flight will be landing.


What is the effect of moving the centre of gravity ahead or behind the aerodynamic centre for HALE UAVs? I have heard a lot about keeping the centre of gravity (CG) behind the aerodynamic centre (AC)will give better performance than keeping CG ahead of AC (for UAVs require high performance). However, I could not find any example UAV/glider/sailplane which operates with CG behind AC. Can some one tell me whether there are any existing UAV/glider/sailplane with CG behind AC position?

Glider pilots in a competition will try to place the center of gravity at the most rearward position possible, sometimes even further back than permitted. This will make the glider statically unstable, but the performance will not be much improved - this is more psychological than real. Don't discount psychology in competitions: Pilots who feel that they fly an inferior plane will take bigger risks to get ahead and have a much higher chance of spoiling their chances by doing something stupid. Therefore, everyone tries to tweak their plane, even if the tweaking only helps to give a good feeling.

Now to the physics: If the center of gravity (CG) is right above or below the aerodynamic center (AC), the aircraft's longitudinal static stability is indifferent. Move it forward and stability increases, move it further back and the plane becomes unstable. The aerodynamic center is where the angle-of-atttack-dependent lift forces act, and in an indifferent configuration the lift coefficient on wing and tail will be equal. Since the aspect ratio of the wing is normally larger than that of the tail, an indifferent configuration will already produce slightly more induced drag for the same lift than a slightly stable configuration. Details depend on the lift loss at the center of the wing due to the fuselage and the combined vorticity of wing and tail, but as a first approximation the above is valid. Moving the CG further aft only makes matters worse.

Now it can be argued that the rear CG location helps in maneuvering the aircraft. This might even be helpful with gliders, since they need to maneuver quickly into the center of a thermal, but HALE (high altitude long endurance) UAVs fly mostly straight, so maneuvers will be few, slow and deliberate.

In essence, moving the CG back in a very stable configuration helps, but the optimum is reached already before the CG reaches the AC. Moving it back even further is detrimental to performance and runs the risk of stalling the tail first due to its high loading.

I do not know of any HALE which flies in a statically unstable configuration, and the glider pilots who do will not openly talk about it.

The only application of static instability (CG behind AC) which boosts performance can be found in supersonic aircraft. And the performance improvement is most pronounced in supersonic flight because it reduces the losses from stability of a highly stable configuration. When incorporated from the start, subsonic instability enables the aircraft to use smaller wing and tail surfaces because both will work less against each other at supersonic speed.


Is line training (pay to fly) allowed in the USA? In Europe some commercial airlines--such as Wizz Air and Air Berlin--charge licensed co-pilots to sit in the cockpit.

It is called "pay to fly" or "Line training".

Is it allowed in the USA ? Is there a FAA regulation regarding that?

It is not prohibited in the US, but I don't think it's currently practiced anywhere.  There used to be a school called Tab Express that did something like this, but the most notorious one was an airline called Gulfstream International Airlines that operated as United Express in Florida.  Gulfstream is now operating as Silver Airways and, as far as I know, has discontinued the practice of having the FO pay to be there.


Restricted Area "popping up" around you -- can it happen, and what do you do about it? What happens if you are flying along through a Restricted Area (lets assume you're in a J-3 so you're VFR/no transponder, but you have a handheld radio to talk to ATC with) and  said restricted area is activated while you're still in it?  Would this be considered an operational failing on ATC's part, or would such a thing be blamed on the using agency (i.e. they ought to ask ATC "is anyone using this?" before they take it "hot")?  Are there procedures you should follow in this scenario, other than "fly out of the restricted airspace the fastest way possible and hope whatever hazard caused the R-area to go hot doesn't hit you?"

Assuming your question is for the US airspace.

Short answer:

When the activation periods cannot be totally anticipated, the description of the area must reference the possibility of activation announced by NOTAM.

The NOTAM notice is normally not shorter than 4 hours before the activation.

While IFR flights will be prevented by ATC to enter or transit an active restricted area, non-controlled flights are under the sole responsibility of the pilot, as regard to flying within a restricted area.

A VFR pilot must take NOTAMs for restricted areas into account when preparing the flight (FSS), and should contact ATC when airborne for confirmation of the activation status.

From the pilot standpoint

You should not enter an inactive restricted area without contacting the controlling agency.

From AIM-3-4-3:

  Penetration of restricted areas without authorization from the using
  or controlling agency may be extremely hazardous to the aircraft and
  its occupants

You are not allowed to enter an active restricted area without prior permission.

From the Federal Register, 14 CFR Part 73, §73.13:

  No person may operate an aircraft within a restricted area between the
  designated altitudes and during the time of designation, unless he has
  the advance permission of (a) The using agency described in §73.15; or
  (b) The controlling agency described in §73.17.

FAA ATC, as the controlling agency, manages restricted area transit or avoidance, based on activation status, but only for IFR flights.

A pilot of a non-controlled flight (e.g. VFR) has the sole and complete responsibility for getting all information about the restrictions that may apply during its flight over a restricted area.

The briefing with FSS should include such research, contacting ATC prior to entering the restricted area seems the minimum.

From controlling / using agencies standpoint

A restricted area is part of the SUA (Special Use Airspace areas), and is jointly managed by Using and Controlling agencies.

Procedures for handling airspace matters, 21-2-4, Times of use:

  a. The times of use indicate the period during which the using agency
  is authorized to schedule and use a SUA area. These times should
  reflect when normal operations are expected to occur. In determining
  the times of use, the proponent should select the minimum period
  needed to meet the using agency's requirements. The goal is to capture
  the majority of the day-to-day activities. When the using agency has a
  requirement for intermittent, less frequent use of the airspace
  (outside the specific published timeperiod), a provision to activate
  the airspace by NOTAM may be stated in the SUA legal description.

What does that say? If some periods of activation cannot be anticipated, the phrasing "by NOTAM" must be used.

  NOTAM activation. Use "By NOTAM" or "Other Times by NOTAM" to
  indicate when a NOTAM must be issued in order to activate the area.

Regarding the notice prior to activation by NOTAM:

  (e). NOTAMs should be issued as far in advance as feasible to ensure
  widest dissemination of the information to airspace users. Normally,
  the minimum advance notice should be at least 4 hours prior to the
  activation time.

  NOTE-  Under no circumstances may SUA be activated by
  a NOTAM unless the words "By NOTAM" or "other times by NOTAM" are
  stated in the area's legal description.

A 4-hour notice is the minimum usually accepted for the publication of the NOTAM prior to activation of the restricted area. I assume the exceptions would be handled on a case by case  basis.

An area without a NOTAM mention cannot be activated outside of the published periods.

Example:

R-2305 Gila Bend, AZ
[...]
Time of designation 0700-2300 local time daily, other times by NOTAM.
Controlling agency FAA, Albuquerque ARTCC.
Using agency U.S. Air Force, 58th Fighter Wing, Luke AFB, AZ.

Related:

How to find out if Restricted Area is “hot” before taking off?
What should you do if you are forced to enter restricted or prohibited airspace?
Can gliders access restricted airspace?


Why does the effective alpha/CL increase at wingtip of swept wing? If anybody could explain this in an intuitive way I'd really appreciate it. I haven't been able to find a clear answer on this. I understand that the increase in effective AoA/CL and the spanwise flow & resultant thickened boundary layer at the tip is responsible for the tip stall (if uncorrected for in the wing design), but I can't wrap my head around the reason why the effective AoA and CL increase in the first place. On a rectangular wing they decrease because of the upwash induced from the vorticies, so why doesn't it have the same effect on a swept wing? Most of the explanations I've found on here are geared for aerodynamicists and not the average joe pilot like me, and the explanations given in most pilot resources are overly simple if an explanation at all.

The book "Handling the Big Jets" by D.P Davies says "Too much sweep produces poor oscillatory stability and a tendency for the tip to stall, causing pitch up" and that's it..

ACE the Technical Pilot Interview by Gary V. Bistrow says "A simple swept and/or tapered wing will stall at the tip first [...] This is so because the outer wing section produces a higher wing loading due to the wing taper, which causes a greater angle of incidence to be experienced to a degree where the airflow stalls at the wingtips. The boundary layer layer spanwise airflow, also a result of sweep, further contributes to the airflow stalling at the wing tips."

This ATPL CBT video I watched says the airflow separation starts at the tips because the weak wingtip vortices allow separation earlier than the inboard part of the wing where the stronger vortices from the wider chord reduce the effective angle of attack... what?!

I have an interview coming up where this question could be asked, and I don't feel like I'd be able to deliver an answer with any sort of confidence at this point.

You are reading the wrong sources.

I've never come across the claim that the lift curve slope of a swept wing increases towards the tip. That might happen in ground effect (when the tips are very close to the ground from the swept wing's incidence), but that effect is very weak, or when the wing tip has noticeably less sweep (crescent wing).

But not on "regular" swept wings in normal operation.

If, however, the intended meaning of your question is to ask why the tips of backward swept wings reach stall conditions earlier than the inner wing: That they do indeed. Read on for an explanation.

What your sources describe is foremost caused by the thickening of the outer boundary layer of backward swept wings. This is compounded by aggressive taper - but this taper effect is very similar in unswept wings, too. Next, when you plot aspect ratio over sweep angle, you will find a boundary above which swept wings develop undesirable stall behavior, so the wing's aspect ratio is another factor (which your sources stay mum about).

To prepare for the interview, I advise you to read some of the answers here, like:

Swept wing layout for maximum lift at transsonic Mach numbers.
Sub- and transsonic flow separation.
Stability of swept wings.
Why the better designs use increased sweep at the tip, precisely to lower the local lift curve slope.

And make sure you follow the links in the linked answers as well if something is unclear.


What's field excitation in an electrical system? "A field relay in the GCU (generator control unit) allows or prevents field excitation within the generator. When open, the field relay deprives the associated power relay of its ground, which causes the power relay to open."

It's an excerpt from the Citation M2's pilot training manual. What's field excitation here?

To my knowledge the M2 uses a Williams FJ-44 engine which has a starter generator on it.  My recollection is that it is a 200A 28VDC starter-generator (SG) which is directly connected to the accessory drive (along with the oil pump).  If 24-28VDC is applied to it, it will start turning the turbine, for starting the engine. At a certain point, once the engine starts, it is switched so that no power goes into it, and instead power is derived from it to charge the battery and provide gen power to the bus.

A typical engine starter configuration is shown above, and is common in smaller turbine engines. Note that the starter generator functions as an electric motor, and also as a power source. (Ignore the circuitry for ignition, throttle relays, etc.)

The SG is NOT an AC generator as is common on transport sized aircraft.  It is also NOT an alternator, as is common to piston engine aircraft.

You may find the descriptions in this manual http://content.aviation-safety-bureau.com/allmembers/faa-h-8083-32-amt-powerplant/sections/chapter-5.php#4. helpful in general understanding of SG operation.  Some of the terms vary from SG vendor to SG vendor, but in general, they are DC motors, and when electric power is removed from them, and they are spinning, they will produce power. Some are self excited with residual magnetism and others utilize an internal circuit to take the small power created by residual magnetism to create a stronger field to create a larger output power from the generator.

Having said all of this, normally a field excitation coil, is the name attributed to an alternator field, which is a magnetic field created by current (not residual magnetism, and not a permanent magnet).  An alternator creates AC internally (not DC like a SG) and is then rectified by diodes which are almost always integral to the alternator. The alternator is usually a three phase device and will have 3 or 6 diodes for rectification. Alternators tend to be less than 150 amps, where SGs are typically 200 to 300 amp devices.

If the SG model is provided, perhaps a refined response is possible.


Does full nose up trim yield best glide in Cessna 172/152 for a given load? I heard from a ground instructor that full nose up trim will yield best glide in a 152 and 172.

As I understand best L/D, a lighter loaded plane will require a slower glide than the same craft loaded more heavily.

Am I correct in thinking that the best glide in the POH is an average glide for an average load?

Was my ground instructor correct? Those extra feet may be useful one day!

That ground instructor is giving you bad information. You can not make a blanket statement like "full nose up trim" and expect it to be valid for any and all situations.

You are correct that the best L/D speed will yield the best glide ratio and that the speed will vary with weight.

The best trim position is the one that will maintain the best glide speed hands off. Any other trim position will require an additional control input to fight the trim and that input will cause additional drag.

So no, full nose up trim will not yield the best glide unless by some remarkable coincidence that is also the trim position for hands off best glide speed.