Unarmed F-16 -- why would it happen? As is documented here among other places, a couple of the interceptors sent after United Flight 93 on September 11, 2001 were scrambled so hastily they did not have time for the ordies to visit and arm them up.  While not having live missiles on you at all times makes sense for a variety of reasons -- I can't understand why a fighter jet would be flying around with 0 rounds of ammo for their gun, considering how much of a W&B impact the ammo load has.

Furthermore, even if the media coverage was referring to them being loaded with training slugs instead of live ammunition -- you can seriously damage an airplane with 20mm slugs, as one Grumman test pilot (among other folks, it seems) found out the very hard way!  Sure, you'd have to use quite a bit of cannon fire to do it, but it'd still be preferable to a last-ditch ramming attack, no?

So, how could a fighter be scrambled into a situation where their only option was to ram their target?

If the article is to be believed, the gun was probably loaded with M55A1 training rounds which are a steel core projectile with an aluminum nose-cap used for target practice as opposed to the M56 HEI or M53 API rounds which the jet would go to war with.  If it had either a full or partial magazine of the M55 ammunition, it would have had at least some firepower available to it, but I imagine the crew had resolved themselves to, if necessary to protect Washington DC, ramming their aircraft into the jetliner.  It just goes to show you how utterly unprepared this nation was for that kind of a ruse.  Its obvious that this was an anomalous and desperate measure by two very brave ANG pilots who responded to a call using whatever resources were available to them.

Post 9/11 most fighter bases in the nation keep at least two birds housed in a Hardened Aircraft Shelter (HAS) in a 'hot' condition, that is, fully fueled, 520 rounds, 2 x AIM-9s, 2 x AIM-120s, with the INS aligned and two pilots keeping watch round the clock for the alert.


How can pilots fly VFR over-the-top in overcast conditions? It's apparently legal for pilots to fly over the top of clouds and fly VFR. However, I don't understand how it's possible to do so, especially since there is no visual reference to rely on to ensure that you are heading in the right direction. So, how exactly does this work, are there any limitations on this and is it possible to be done safely?

Perhaps the most important rule for a VFR pilot is "see and avoid" -- be able to see any immediate dangers (other aircraft, buildings, ground) and avoid them. When flying above the clouds, you can certainly see other aircraft and avoid them (as long as you're maintaining required cloud clearances), and the same for buildings. You cannot see the ground but on the presumption that your altitude is well above the ground and not significantly changing, this is not a factor.

As long as you're able to see the horizon, you're unlikely to become disoriented to the point of not actually knowing that the aircraft is or isn't flying straight and level. (Contrast that with flying in the clouds, where it's easy to become spatially disoriented and not be aware that you're starting to dive.) Limitations are basic: it's the PIC's responsibility to follow the regulations.

As to heading in the right direction and navigation in general, VFR flight does not mean one cannot rely on radio aids to navigation. VOR and GPS certainly make navigation very possible while on top, but to get to your question of "can it be done safely?" -- I suppose this comes down to a question of how reliable are your electronics and where are the weak spots; is the pilot under flight following or otherwise in contact with ATC, etc. If one were on top and were to lose the electrical system, there's certainly an emergency to be declared. If already in contact with ATC, it's an emergency that can likely be managed with the utmost of safety.


How to calculate the Pitching moment coefficient of wing? I want to check the trim condition for aircraft. For this I need the pitching moment coefficient of wing. How to calculate?

Take the pitching moment coefficient of the wing's airfoil at the trim angle of attack. Since the reference area normally is the wing's area and the reference chord is the wing's mean aerodynamic chord, you do not even need to adjust the number. I hope you have the airfoil polar at hand; if the wing uses several airfoils, take a volume-weighted average; volume being here the product of local area and chord. If the wing has a symmetric airfoil, the coefficient is zero, in most other cases it will be a small negative number.


Why are the transition altitudes/levels different around the World? I got tasked by my instructor pilot to find out why each country has a different transition altitudes / levels. For example in the US it's FL180 and in Saudi Arabia it's FL150.

Can someone help me with that, and if possible provide any document that talks in details about this subject.

Flight Levels are easier to work with for ATC and for pilots since everybody will use standard pressure setting. Effectively it means less work load and lower probability of making an error with a safety impact.

Typically the highest terrain elevation plus some margin is used as transition altitude. This gives the highest number of flight levels available.

For example in the Netherlands, being a flat country partly below sea level and the highest obstactle at about 1500 ft, has a transition altitude of 3000 ft for IFR flights and 3500 ft for VFR flights.

There is (was?) an initiative to create a common transition in Europe, but I haven't heard anything about it for a while. Some interesting reference documents were produced:

Towards a Common Transition Altitude - A Flight Deck Perspective

A Common European Transition Altitude  - An ATC Perspective

(both files in PDF format)


Were cabin lights really consuming that much power back in 2006? Even though it is nine years since then, I can still clearly recall my return flight from London Luton to Katowice (Poland), that I took on July 31, 2006 evening. There were a huge, strong storm front coming to Poland (ten days of extremely high warmth followed by ten days of an unusually high raining and really strong stroms, that started on August 1, 2006 in Poland), that we hit somewhere over Germany (hit or flown somewhere around / near it).

I was then and I am now a layman in aviation, so my description and terms used my not be correct, however I recall, that it was really a bumpy ride. Plane was falling down (loosing 100 ft or so) and trying to  go back to previous altitude etc.

At some point, one of a pilots informed everyone on board, that they'll be dimming all the cabin lights, because "they need more power to the engines to go through this strom". And soon after that, they really did that and for about 10-20 minutes we were flying in a very limited cabin lightning.

Is this true, what we have heard nine years ago? Were cabin lights really consuming that much power back in 2006? Or was that just a strange way pilots wanted us to get a little bit more scared?

I was flying with Wizzair, so I assume, that it was A320. There is a similar answer on similar matter, but it says about LED-based cabin lights system and I can hardly believe, they were introduced to "cheap" Wizzair's A320s back in 2006.

But, even assuming, that our plane was fitted with a "typical" lights, much more energy consuming that LEDs, it is really hard for me to believe, that cutting on cabin lights could bring any noticeable power to airplane's engines. Correct me, if I'm wrong.

Thunderstorms normally mean icing conditions, which necessitates anti-ice, both through bleed air and electrical systems.

I can see this going a few ways:

Pilot is (for instance) at high altitude and really does need the additional bit of power to cross above the thunderstorm.
Pilot has flown aircraft type where this is a legitimate problem and figures 'why not?' for safety sake.
Could be something about motion sickness in the storm, since closing your eyes can be an effective countermeasure. Perhaps dimming the lights goes along the same lines.


Is there an Instrument Flight Procedures database in digital format such as ARINC 424 for Germany? I am aware that Eurocontrol has an online repository for charts related to Europe's Instrument Flight Procedures. However, I am trying to find these data coded in a digital format such as ARINC 424 or something similar, especially for Germany. The ideal situation would be something similar to FAA's Coded Instrument Flight Procedures.

Could someone point me in the right direction? I also tried to check Jeppesen's website (link) but I keep receiving a "Page not found" error.

Eurocontrol does offer the whole thing as well – albeit not for free as the FAA does. Check out https://www.eurocontrol.int/service/european-ais-database for more information.


What effect does cold weather have on helicopter performance? Recently I was flying around my RC helicopter on a 40 $^\circ$F (4$^\circ$C).

At some point in the flight the rotor just stopped moving in mid-air and then I regained control back at ground level. I never had any issues before in warmer weather, so I assume it's a cold-weather related issue with this model.

This leaves me with two questions:

Could something like this  happen in a full-size helicopter like a Robinson R22?
What effect does temperature have on helicopter performance in general?

Cold weather definitely affects the performance of helicopters (as it does all aircraft), because it affects the density of the air they move through.

I will ignore icing so let's assume dry air.

The lift equation is:

$$ L = \frac{1}{2} \rho V^2 S C_L $$

$L$ = Lift

$\rho$ = density of the air.
$V$ = velocity of the aerofoil (rotor)
$S$ = the wing area of the rotor
$C_L$ = Coefficient of lift , which is determined by the type of rotor and angle of attack.

As you can see, lift is proportional to the density of the air and since density at any given altitude is inversely proportional to the temperature, the colder the air, the more dense it is.  So, for a given rotor, spinning at a given speed at a given altitude, the colder the air, the more lift is produced.

Engines also operate more efficiently when the air is colder since the density is higher and the charge of air/fuel introduced into the engine is greater. (See here for more information.
 Of course, this does not apply to your RCH apart from some effects on the efficiency of the batteries and motors (which may perform more poorly at lower temperatures).

So in general, the colder it is, the higher the performance.  In most helicopters, the difference in performance on a cold, dry winter day is very noticeable and much better than on a humid, hot summer day.

It's also worth understanding the properties of air.  The colder the air, the lower the dewpoint so the less moisture it can contain.  Moisture in the air also lowers peformance.

The problems you encountered with your RC model aren't problems with helicopters in general.

Colder weather does not make you more susceptible to vortex ring state (aside from some minor side effects): You would still only descend into the vortex if you did not pilot the helicopter correctly. (Note that you cannot enter vortex ring state with the rotors stopped since they are not producing a vortex.  Also note that on a real helicopter, if the rotors ever stop, they will not start again no matter what you do.)

BTW, I'm being pedantic, but helicopters do not have propellors.  Technically, they are the same thing, but asking a helicopter pilot about the propellor annoys them, even if they don't show it.  It's as bad a calling a helicopter a "chopper".  Only Arnie and Chuck get to call helicopters "choppers".  And that's only because helicopter pilots are universally weedy and would get stomped by Arnie or Chuck if they argued.


Is there any book like the "Stick and Rudder" about the helicopter? I have read the Stick and Rudder, that's really a classic book. Now I want to study the flight theory about helicopters, is there any book like this?

There's a book called "Cyclic and Collective" by Shawn Coyle. It's a good book, although it is a little bit more detailed than something like "Stick and Rudder". There's a lot of good stuff in there, though.


Why is there a grid below the antennas of an VOR? A few days ago I made a trip to the VOR-DME near my local airport LOWG, because I always wanted to see one. It's GRZ on 116.2.

When I was there, I wondered why there is a giant metal grid under the actual antennas of the VOR.

Source: Klaus Krasser

Close picture:

Source: Klaus Krasser

The grids squares have a size of estimated 10x10 centimeters.

What is this grid for? Is it maybe used as a shield or reflector? 116.2 Mhz has a wavelength of ~2.6 meters, so the grid would reflect the radiation. If thats the use of the grid, why is it needed?

I think this is an example of a counterpoise:

  [It] is a network of suspended horizontal wires or cables (or a metal screen), used as a substitute for an earth (ground) connection in a radio antenna system (...) when a normal earth ground cannot be used because of high soil resistance or when an antenna is mounted above ground level.

Generally used for LF and VLF antennas as that is where they are most efficient.

An example from that Wikipedia page of an AM radio antenna counterpoise:


Which type of airfoil does the 737 use? NACA airfoils are not used anymore except on some aircraft. So which type of airfoil does the 737 use? Which company designs it and builds it?

Boeing designs its own airfoils for its aricraft. According to this site, the following airfoils are used in 737:

Boeing 737-100: Wing root- BAC 449/450/451; Wing tip- BAC 442

Boeing 737-200: Wing root- BAC 449/450/451; Wing tip- BAC 442

Boeing 737-300: Wing root- BAC 449/450/451; Wing tip- BAC 442 mod

Boeing 737-400: Wing root- BAC 449/450/451; Wing tip- BAC 442 mod

Boeing 737-500: Wing root- BAC 449/450/451; Wing tip- BAC 442 mod

The 737 wings are manufactured at the Boeing plant in Renton, Washington.


How does turbojet thrust change with altitude? I am studying the rate of climb of turbojets, I don't know how the thrust varies with altitude and in particular which are the differences between necessary thrust and effective thrust.
Hope being clear, I can't translate properly maybe so ask and edit !

The thrust variation with altitude would be highly engine specific, but the general trend is nicely depicted in the image below:

Read this for further details.


If a typical passenger plane had total failure of all engines mid-flight, is it possible for passengers to survive? I have pretty bad flying phobia and often wonder what would happen if all the engines on the craft went out at the same time. I initially had the anxiety-ridden thought that the hundreds of thousands of pounds of metal would nose-dive and explode into the Earth. Then I read somewhere that the plane would actually glide forward as it descended, making a landing where all passengers survive a possibility. What if you were flying over the ocean versus flat land versus mountains? Has complete engine failure ever happened?

Air Transat Flight 236 experienced a complete power loss over the Atlantic Ocean in 2001. Yes, all passengers and crew survived after the aircraft glided 75 miles to a runway on the Azores islands.

Even in the event of the loss of all engines, an aircraft can keep its critical electrical systems running thanks to the ram air turbine which allows the crew to maintain control of the aircraft and to communicate with ground crews to determine the best course of action for the situation. This was the case with Air Transat Flight 236.

Moreover, consider US Airways Flight 1549 which experienced a complete power loss after takeoff due to a double bird strike. Even in the situation where the pilot had very few options due to being at such a low altitude and being over a populated area, all passengers and crew survived after ditching in the Hudson River. However, in all fairness, ditching attempts do not normally end well.

In all, complete power loss in a modern airliner is extraordinarily rare. While gathering links for this answer, I ran across the "List of airline flights that required gliding" article on Wikipedia. The short length of the list speaks for itself (4 flights in the last decade).


Is an F-16 able to fire any of its weapons while on the ground? I can recall at least one Hollywood movie, where this takes place (although not an F-16 in the movie).

Is this at all possible? Can the F-16 fire its M61 while taxiing on the ground? What about other weapons?

Note that I'm not even sure myself what good this would do or what the point would be. I'm just curious.

I can imagine the bullets would not even be able to hit a target on the ground, perhaps pass over or below the target. Does the M61 even point perfectly parallel to the ground on an F-16? At least it would change its elevation by some degrees up and down when taxiing, braking, etc. I suppose I could imagine helmet cuing of a low-flying target with an AMRAAM while the aircraft is still trying to take off. Since it has happened before that an airbase is taken by surprise, perhaps this question is not as silly as I at first thought. However, I'm under no illusion that an AMRAAM would be able to engage any target when fired from an aircraft on the ground---doesn't the missile need some time and space to drop from the rail and fire its engine?

I'd love to be told that I'm wrong though.

Nevertheless, I'm curious if the aircraft is designed to prevent firing its weapons while the landing gear is out, or a similar safety feature. If so, can it be overridden? Specifically, what, if anything, prevents an F-16 from firing its M61 while on the ground?

As a bonus question, has this by accident ever occurred, that an aircraft (F-16, or other) fired its weapons while stationary, or taxiing around at an airport?

Yes, the F-16 can fire on the ground, and it has happened as an accident

A maintenance worker accidentally fired a 20mm Vulcan cannon from an F-16 jet he was working on at Belgium's Florennes Air Force Base earlier this week, destroying another F-16 while damaging another aircraft nearby, according to multiple media reports.
A worker apparently opened fire across the flight line while performing maintenance on Oct. 11, according to Avio News. The F-16 that was racked by cannon fire quickly caught fire and exploded, since it was loaded with fuel and being readied for an afternoon sortie, according to Scramble Magazine.

https://taskandpurpose.com/f16-destroyed-belgium


What would cause an aircraft (j3 piper cub) to slide backwards when going into a stall? Fifty years ago, I took lessons at an uncontrolled airport and this happened to me, fortunately my instructor took over and as cool as a cucumber corrected the situation. His detailing of what he was doing fell on deaf ears.

This is called a "tail slide" or "whip stall". It is an aerobatic maneuver when performed on purpose, but can be dangerous if done by accident because it causes damaging forces to control surfaces. Some aircraft if the CG is in the wrong location, will tend to tail slide before they fully stall (my C-177 is prone to this). – Ron Beyer

Aerobatically, to enter a tail slide you would enter from level flight at full speed (maneuvering speed) and pull straight up (1/4 loop) and wait for your airspeed to bleed off to zero. The aircraft will then slide backwards and eventually the nose will drop. You stay in the dive until you have enough airspeed to recover.

Doing this on accident is a little more complicated, and is surprising that the J-3 could even do it. What happens is that you pull up while adding power just as you would be practicing a power-on stall, only the tail stalls before the wings do. The tail drops and the aircraft slides backwards. This is especially dangerous at low altitude because the only way to recover from a full tail-slide is to allow the nose to drop and gain airspeed in the dive.

The recovery action for an incipient tail-slide is similar to recovering from a stall. Nose down to the horizon and full power. Newer training requirements don't let the student go into a full stall, it teaches to recover the aircraft as soon as the aircraft begins to buffet, or at stall speed + 10kts. I'm not sure I agree with the new requirements but that is a topic for another question.


Has the FAA ever granted a LODA for compensated instruction in a non-LSA E-AB? I have been trying to find a list of granted LODAs to get this question answered. No joy. I am considering purchasing a Sling TSI E-AB. It would be great to be able to instruct in it for compensation.

You may have to contact the FAA directly to get an answer on that but it seem like an unlikely case. Since you are looking for a LODA that covers "training flights" you would need a LODA SFA

  D.    LODA Special Flight Authorization (LODA SFA). A Flight Standards
  office normally issues a LODA in cases where frequent operations of a
  specific nature occur over a protracted period of time. However, there
  may be situations such as ferry flights, sales demonstration flights,
  and training flights where the 60-day requirement may not be feasible.
  For these situations, a Flight Standards office may issue a LODA SFA
  on a case-by-case basis for specific flights. A LODA SFA is a LODA
  authorizing an operator to accomplish short-term operations (normally
  a maximum of 30 days) in accordance with 14 CFR part 91, and can be
  issued for the following:

LODA SFA's are short term as discussed later on, specially they are typically 30 days in duration

  ...

  2)    A LODA SFA is conducted in accordance with part 91 rules. The maximum duration of a LODA SFA is normally 30 days.

  ...

To regularly provide instruction in the aircraft you mention you would effectively have to continuously apply for a LODA SFA on a rolling basis, something the FAA likely will stop approving eventually. I would say that your use case does not really fall under the broader spirit of the regulation

  ...

  LODAs to all of part 125 are issued for operational uniqueness

  ...

Although you could make a case if you were offering training to others specifically looking to buy that aircraft. If you are just looking to offer flight instruction the FAA is likely to tell you to go find an approved airframe and play by the rules. I cant find an exact price for a sling kit but Im seeing north of 100K for built ones, even if the kit costs half of that @ 50K you can get a whole lot of 172 or Cherokee 160 for that price and offer instruction without the hassle of special authorization (above whats normally needed).


How much air, by mass, enters an average CFM56 turbofan engine cruising per minute? How many kg of air enters an average CFM56 turbofan engine cruising per minute?

I have looked everywhere and cannot find it or figure out how to find it, if you could either tell me or give advice on where or how to find it I would appreciate it. Thanks

The CFM56-5C, the largest CFM56, powers A340. Its thrust is about 145 kN at takeoff. It has an air mass flow of 485 kg/s (source).

CFM56-5C, source.

The Trent XWB (A350 XWB), a larger engine developing 430 kN at takeoff, has an air mass flow of 1,440 kg/s (source).

T/O vs cruise power and fuel consumption

According to this document, the CFM56-5C T/O thrust is 150 kN and the cruise thrust is 40 kN (other sources states 30 kN), 27% of the T/O thrust. Related fuel consumption are 2.3 and 0.5 kg/s, meaning cruise fuel consumption is 22% of the T/O fuel consumption.

Air mass vs fuel mass

For a fuel consumption of 22% of the T/O consumption, it can be assumed relatively safely that the air flow is in proportion, so we can roughly estimate air mass flow in cruise to be about 485 * 0.22 = 107 kg/s. This is supported by figures found online in non authoritative sources, like page 245 of this book:


Why is this 777 with doors and windows covered? It's in the desert somewhere and it has its windows and landing gear and doors covered. Why is it that? Of course, it's not in service anymore...
Source

The aircraft is in storage. The Boeing 777-212(ER) seems to operated by Singapore Airlines under lease from Pembroke before being withdrawn from use in Dec 2014. After that, the aircraft is under storage at Southern California Logistics Airport.

The windows are covered in aluminium foil for storage; however, the engines appear to be still there (which will usually be removed), which indicates that the storage was planned for only a short time.


Does a biplane need wing twist? Does a biplane need wing twist?

I understand wing twist is used to:

creat a soft/partial stall ( hopefully create a partial stall before a full stall)
balance the wing so a smaller horizontal tail is needed with less drag.

So if the top wing on a biplane is canted upward, in a mild stall, it will stall first, with the bottom wing still flying, creating a partial stall.

If the bottom is at a lower AOA, it's like having wing twist on 50% of the plane. Wouldn't this act like wing twist on a monoplane and reduce the total moment, therefore needing less of a horizontal tail, like the self balancing wings on modern gliders. Is this correct, at least conceptually?

Evidently, they don't. The picture shows a fully aerobatic Pitts Special from 1966 with no visible wing twist.


What factors besides tail/headwinds cause the time taken to fly some routes to be different from the time it takes to return? I've observed this on many flight trackers; the time taken to fly from A to B is not necessarily the same as the time taken to fly back from B to A. Why is this so and what external factors besides prevailing winds (headwind going one way, tailwind the other) create this difference in time taken?

Example: late May 2015, per American Airlines, PHL to MCO is about 15 minutes longer going south than going north, but times MIA to BDL are essentially identical in either direction.  Prevailing winds would affect both routes similarly, but the difference in times northbound vs southbound isn't the same.  Why not?

Sometimes, prevailing winds will account for the difference -- flights going west to east in the US will generally be scheduled for less time than flights going the opposite direction.

The other major factor that can play is in taxi times.  If the delays waiting to get in to a gate matched the times waiting to depart at each airport, these delays would affect both legs equally, but that's not generally the case.  Most of the time, at a congested airport like Atlanta, you can generally get into the gate fairly quickly, but the delays getting airborne can be far longer.

Thus, a flight scheduled from Podunk into ATL might be scheduled for 10 minutes taxi from the gate to departure at Podunk, the flight time, and another 10 or 15 minutes to taxi in, but in the other direction they might build in 30-45 minutes of taxi time departing ATL, plus the flight time, plus 5 minutes landing until blocking in at Podunk.  Thus the leg into ATL is scheduled for flight time plus 25 minutes, while the leg from ATL is scheduled for flight time plus 35-50 minutes.

If it's winter time and your flight was going from BOS or PHL to MCO, another factor in addition to the potential for massive departure delays would be the need to build in time for deicing -- done before departure in PHL or BOS, but not so much leaving Orlando.

Of course, if you didn't need to deice "today," then you'd probably beat the scheduled time by a good bit.  Likewise if the weather was clear and the airport had no delays.  But the mega-hub airports run so close to capacity that any little hiccup in the weather or otherwise causes delays to start building, and then you're using all of your scheduled time, and sometimes more.


What is the relation between maximum airspeed and engine power? I know that the drag force on an airplane scales as (velocity)^2 which implies that doubling the airspeed will produce a 4x increase in the drag force. So my question is:

Imagine we have an airplane that flies 100MPH on 100HP. If we replace the engine with one that produces twice the power (200HP) I would anticipate that, everything else being equal, the most speed increase I could possibly expect would be (100MPH) x sqrt(2) or about 140MPH- and that if I wanted to double the airspeed I would need a 400HP engine. Is this correct?

In theory, yes. However, everything else isn't equal.

When you double the speed, you probably exceed the design parameters of the airframe. You'll definitely exceed the VNE of a 100mph airframe if you operate it at 200mph, and then all sorts of ugly things can start happening, from weird handling characteristics to large pieces, like the tail, coming off inflight. Remember that 4x drag means 4x the strain on the airframe. Is your airframe up to the task?

Most 100mph aircraft have a design optimized for short takeoff and landing, wings that are very efficient at slow speed, but not efficient at higher speeds. So the simple drag calculation may not hold true, when you take wings designed for slow speed and operate them at a much higher speed. The actual drag will probably be higher, if you consider the design of the airframe.

Compare a 100mph aircraft, say a Cessna 150, to a Mooney 201, that can go 200mph. The Mooney has much thinner wings, that result in a high landing speed, but lower drag at higher speeds, while the 150 has a fatter wing that gives a lower landing speed and better low speed handling, but may not fare so well at a much higher speed than it was designed to handle. The other tradeoff, aside from speed, is the 150 is far more forgiving of a clumsy pilot, while the Mooney is probably not the best trainer for first time students.

Note that the Cessna goes 100mph on 100hp, while the Mooney goes 200mph on 200hp, so the advantages of designing for the target speed can be seen in not just upping the power, but redesigning the airframe (and adding retractable gear).


What is this airplane with small wings at different angles seen at Paphos Airport? I landed in Paphos Airport in Cyprus and spotted this thing sitting in the sun near the runway. It has small 'wings' sticking out at different angles. What kind of plane is this? (Sorry for the bad photo.)

Looks like a QAC Quickie Q2. It's a kit plane with over 2,000 produced.

(wikimedia.org)


How do modern helicopters tackle Vortex Ring State? According to the wikipedia page :

  Air vortices can form around the main rotor of a helicopter, causing a dangerous condition known as vortex ring state (VRS) or "settling with power". In this condition, air that moves down through the rotor turns outward, then up, inward, and then down through the rotor again. This re-circulation of flow can negate much of the lifting force and cause a catastrophic loss of altitude. Applying more power (increasing collective pitch) serves to further accelerate the down wash through which the main-rotor is descending, exacerbating the condition.

The vortex ring state has also been listed as a potential hazard in the Wikipedia page for helicopters. What are the various preventive measures used to prevent such a condition?

Helicopter driver here.

I am not aware of any technical features which can prevent it.  Quite simply, the disc must accelerate air downwards to generate lift.  If the helicopter descends into that downward accelerating air, it is recycled through the disc resulting in a catastrophic loss of lift.

The preventative measures are:

Be aware of the flight conditions under which vortex ring state may occur.
Do not fly into these conditions.
Be able to recongise the symptoms of vortex ring state.
Understand, and be practised in, the correct recovery technique, which is most often fly in any direction away from the vortex.

I am not an aerodynamics engineer but I doubt that there is a physical solution to this - and if there was, it would likely have been implemented by now.


What type of engines are used on the RNLAF KDC-10 and KC-10's? Does anyone know what type of engines are used on the Royal Netherlands Air Force's KC-10's/KDC-10's? I did some research and found out that they are GE CF6-50 engines, but what comes after the "50"? Is it a -50C2B, for example? Or a -50A?

The Dutch defence department website has the answer:

  motoren: 3 maal CF6-50C2 turbofan

Which translates to

  engines: 3 times CF6-50C2 turbofan


How much power does the APU generate in a Boeing 737-300? I have just watched Mayday episode 11, season 11, "Nowhere to Land" about TACA Flight 110 incident and I was more than surprised to learn that:

APU must be turned on,
APU needs about 5-10 minutes to start powering anything,
APU can be used to power lights in the passenger cabin (???).

This is the exact flow of events (plus voice description) as it is shown in above cited episode:

a pilot is shown turning APU's switch,
narrator explains that a certain amount of time must pass before APU starts working,
lights in passenger section are turned on after this period of time passes.

Contrary to the above, I was always told (based on Air Transat Flight 236 incident, shown in season 1, episode 6, "Flying on Empty" of Mayday, and based on other sources):

APU is deployed automatically (gravitational solution) whenever power from engines is lost,
APU starts powering equipment nearly immediately,
APU is able to power basic instruments only, even pilots cabin remains dark, not to mentioning passenger sections.

What is the reality?

You are getting the APU mixed up with the RAT:

The Auxiliary Power Unit (APU) is an engine powered generator, either a gas turbine or a piston engine. APUs deliver enough power to start engines, power cabin lights,  cockpit instruments and radios, and in some cases power hydraulics. Often the APU can only be run at lower altitudes where the air is thicker
A Ram Air Turbine (RAT) is a small, automatically deployed propellor run generator which is lowered into the airflow if there is a loss of power. The RAT generates enough power for the most critical systems only, and is designed to give pilots enough electricity to glide until the APU can be started

Often the APU runs from the same fuel supply as the engines, so if the airplane runs out of fuel the APU is unavailable, so the RAT is all they'd have.


Can military personnel land their own private aircraft at military bases? First off, I am asking similar, but logistically different question to What process do I follow to land a civilian aircraft on a military base?

If I am in the US Military (with a valid ID), am I free to land my personal (civilian style) airplane (with my civilian licence) at military bases - in the same sense that I could drive my car on?

As someone already said, 32 CFR 855, Table 1 states:

  Service members, operating their own aircraft, leased aircraft, or other available aircraft for official duty travel (temporary duty, permanent change of station, etc.) or for private, non revenue flights

This of course means you still need to contact base ops for a Prior Permission Required (PPR) of some sort (if required).  If the base has an Aero club, then your logisitcal considerations are much less (most aircraft probably don't want to refuel on JP-8).  The list of Aero Clubs seems to have decreased since I was in though.

The Flight Infomration Handbook will have all the base requirements listed in it.


What are the effects of an increase in altitude upon ESHP of a turboprop engine? This is one of the questions in the ATP material:

  What effect will an increase in altitude have upon the available
  equivalent shaft horsepower (ESPH) of a turboprop engine?

  A) Lower air density and engine flow will cause a decrease in power
  B) Higher propeller efficiency will cause an increase in usable
  power (ESHP) and thrust
  C) Power will remain the same but propeller
  efficiency will decrease

I chose (C) but the correct answer was (A). Could anyone explain why? Isn't the whole purpose of a turbocharged engine to sustain engine power at high altitude (below critical altitude)? Why does lower density at high altitude cause a decrease in power?

You have confused a turboprop engine with a turbocharged engine. They are very different and have very little in common.

A turboprop engine is powered by a turbine engine geared to a propeller. Turbine engines have less power available at higher altitudes so answer "A" is correct.

A turbocharged engine is a piston engine which is fitted with a small compressor driven by exhaust gas pressure. The compressor (turbocharger) can increase manifold pressure to restore sea level power at altitude. (Turbo-normalized) A turbo charger can also supply greater than sea level manifold pressure for more power. (turbo-supercharged)

Both types of turbocharged piston engines can maintain full rated power as they climb, but they will eventually reach a limit where full power can no longer be maintained.


What is that decommissioned plane near Nadirgul Airfield? This is a decommissioned plane that I found at my place. I'm guessing it is a Dakota DC-47.

The plane is placed near a small airstrip called Nadirgul (name of the location), which lies in the outskirts of the city Hyderabad, India

This is right, this model is a C-47A-85-DL Skytrain, a variant of the "Dak" DC-3 selected by the US Army Air Force in 1941 as its standard transport aircraft.

Image from Le Bourget Air & Space Museum

It has been early used by Allied forces to drop paratroopers and equipment or to tow gliders behind the coastline for the D-Day and operation Overlord.

C47 from 439th Troop Carrier Group and Waco with their invasion stripes taking off from England for Utah Beach, June 1944 (source)

It had a long history and was everywhere until the end of the war and in the post-war conflicts, e.g. many of them participated to the Berlin Airlift:

C47s unloading at Templehof airport in Berlin, 1948. Source

Differences with the DC-3 include:

cargo door
hoist attachment
strengthened floor
shortened tail cone for glider-towing shackles
astrodome in the cabin roof.

You said the aircraft was standing at Nadargul airstrip. This information has been very useful to find the exact aircraft. It happens the aircraft did have a long story, throughout eras of war and peace, linking people of many countries:

A C-47A-85-DL airframe with c/n 20012, is manufactured in Long Beach, California, by Douglas (source) for the USAF (registration 43-15546).
In 1946, right after the war, it becomes the French government property (F-BAXK) operated for cargo flights under Air France colors.
Sold in 1950 to Portuguese government to be operated from Lisbon by Aéro Portuguesa (CS-ADB).
1953: Leaving Europe when sold to Lebanon, operated by Air Liban (OD-ABQ).
Finally moving further eastwards to India, bought by Kasturi and Sons (VT-DTS) to deliver The Indu newspaper in distant areas until it is decommissioned and transferred to Nadargul where it is used now by the Flytech Aviation Academy for ground training (source).

(Source)

A movie star

This aircraft is seen in the 1965 Hindi feature film, Gumnaam (video):

Stamps in Israel

The aircraft was used for the first aerial mail sacks sent in 1948 from Israel to European countries, when it belonged to Air France. Two stamps commemorated in 2008 "60 Years of Friendship between Israel and France":

  On June 9, 1948 an Air France DC3 airplane, registration number
  F-BAXK, landed Haifa, flown by pilot Georges Bernard and two crew
  members: communication technician Mr. Sortais and mechanic Mr. Boutet.
  The plane made an interim stop in Bucharest while en route to Israel,
  thus arriving with mail from both France and Romania. The plane
  subsequently left Haifa on June 11 at 6:36 a.m., bearing mail to
  France and other countries with which Israel had re-established postal
  ties: Bulgaria, Holland, Hungary, Yugoslavia, Poland and Switzerland.
  It also carried a mail sack bound for New York. After interim stop in
  Cyprus, Greece Italy and Nice in southern France, the airplane took of
  the next day towards its final destination, Le Bourget airport in
  Paris, where it landed on June 12, 1948 at 11:01 a.m. and delivered its
  mail sacks.

(Source)


Why would a C150 tachometer be sluggish? After having an issue with a tachometer bouncing so much that it fell off, I now have the opposite problem with the replacement (Mitchell D1-112-5025): The tachometer is sluggish. When I open the throttle from 1,000 to 1,700 for the runup, the tachometer takes about 6s to show 1,700 RPM. The delay seems even more pronounced when I reduce power. The smaller the change in RPM, the longer it takes the tachometer to update.

My A&P says that it just needs breaking in, which could possible because the sluggishness has reduced since the tachometer was installed; however, I was under the impression that modern tachometers should work without sluggishness immediately.

The counter increments as expected.

How should I approach this problem? Is it common for tachometers to need >15 hrs to break in? Is it possible for a drive cable to cause this issue (continue to spin, but in such a way that the tach speed is sluggish)? Could some kind of installation issue have caused this problem?

UPDATE Mar-22: Removed the tach cable, inspected outer and inner cable, washed, dried, lubed with graphite, and replaced. Cable seemed good, but the tachometer is still sluggish.

UPDATE Mar-24: Tachometer is now beginning to exhibit the same symptoms as the previous tachometer before the needle spun off; the tach is still sluggish, but the needle is now bouncing considerably.

No you don't have to "break in" flight instruments.

When a TSO'ed instrument is new it will get an Acceptance Test Procedure (a functional test) from the factory done before it is released and it has to perform to specification.  There is no such thing as a break in in service.

That being said, sometimes parts get by the ATP and malfunction in service, so you never know.  In my years in the regional airline industry, "bad from stock" is not uncommon even with super fancy Transport Category components.

Problems with the flex drive could be a factor, but a flex shaft would have to be binding and winding up a hell of a lot to create that kind of lag.  It's still something that could be checked.

Before that though, you should call Mitchell's tech support and describe what it's doing and see what they say.   Maybe they'll tell you there is a break in, but I'd be pretty surprised if they did.


What is this airplane I saw with propellers behind the engine and canards on the top of the plane? While playing Pokémon GO, I saw a plane with:

Propellers behind the engine.
Canards at the tip of the plane.
Normal-looking tail with a flat wing at the end of the upright bit.

I know these are probably not the correct terms, but I'm not sure what else to call those components.

The "propellers behind the engine" is called a "push configuration". One you look through aircraft that are private business planes, the Piaggio P.180 Avanti comes up as a likely contender.


Why did Britain develop three different bombers (V-bombers) after WW2 instead of one? Why did Great Britain develop three different planes - Victor, Valiant and Vulcan - after WW2 for the same task?

I suppose that at that time the US and the Soviet Union were producing high numbers of bombers too, but did those countries stick to one design? Why did France not produce different bombers? My main question is about the effort of Great Britain and I am listing other countries only as examples.

Here is the Handley Page Victor:
(source)

Here is the Vickers Valiant:
(source)

and the Avro Vulcan:
(source)

I've heard numerous reasons cited - perhaps due to the secrecy around all things nuclear - but from what I know, there are a few reasons:

Speed of deployment

Of the three designs, the Valiant was simplest, and could therefore be built and deployed the soonest. The initial Valiant prototype didn't actually meet the requirements that had been set forth for a nuclear-capable bomber aircraft, but it was better than nothing and with the mounting threat that the Cold War might turn hot, this speed was an asset, so the requirements were changed to allow the Valiant to enter service. This would fill the gap in the UK's forces, and more advanced aircraft could follow later.

Specialisation

Each of the V-bombers had slightly different performance, meaning that they could perform slightly different roles.

The Victor had the greatest payload capacity (up to 35 thousand-pound bombs)
The Vulcan was smaller and more agile and could therefore use smaller runways, and could just about break Mach 1 in a slight dive. It could also be refuelled in flight.
The Valiant was the simplest design, and therefore the most reliable (although it later emerged that a newly-developed and poorly-understood alloy introduced some serious design flaws)

Spreading the risk

Overall, the differences between the various aircraft were pretty minor, and the main reason that three contracts were allowed to progress to delivery was simple: the UK didn't want to put all their eggs in one aeroplane. The need for a nuclear deterrent - and a means to deliver it - was so great that it could not be jeopardised by a poorly-designed bomber, a mismanaged project, unforeseen failure of the company building the aircraft, etc. By having three aircraft from three different manufacturers, the UK greatly improved their chances of having at least one aircraft that actually worked, was delivered on time, and was able to remain in service for a reasonable period of time.


Would tubofan aircraft benefit from adjustable fan blades? Would turbofan aircraft benefit from adjustable fan blades as do turboprops? Propeller aircraft adjust their propeller size, i.e. shorter propeller for cruising at maximum design altitudes where the air is thinner for efficiency, and a longer propeller where the air denser for increased thrust used for increasing altitude at early stages of flight. Would turbofans benefit from this when flying at maximum altitude for efficiency? If not why not?

Both turboprops and turbofans want their blades to be as long as they can be, in all circumstances: that way the largest amount of air is accelerated, and a given amount of thrust created with the highest efficiency. It never pays to reduce the bypass ratio of a subsonic turbofan.

Turbofans with adjustable blade pitch would make sense though, like the constant speed propellers that turboprop engines already have. This article from Flight Magazine 1973 gives the advantages of a high bypass variable pitch turbofan:

  Inherently low noise levels because of high bypass ratio and low fan
  tip speeds.
  Rapid thrust response, allowing thrust modulation for
  accurate rate of descent control on the approach and maximum use of
  reverse thrust during the landing run.
  Reverse thrust available
  without the weight or maintenance penalty associated with reversers
  and spoilers.
  High ratio of take-off to cruise thrust well matched to
  the needs of short-range, Stol aircraft.
  Good cycle efficiency and specific fuel consumption.
  Wide engine operating margins.
  Provides many of the advantages of the turboprop while avoiding vibration,
  ground and fuselage clearance problems and allowing better matching
  with higher cruise speeds.


Can a captain fly in the right seat as SIC under part 129? Re a Part 129 air carrier that is short of co-pilots:  Can a pilot who usually flies in the left seat (B737 aircraft: non-US registry) as the Pilot-in-Command fly in the right seat as the SIC (co-pilot)?

From the regulations:

  Each person acting as a flightcrew member must hold a certificate or
  license that shows the person's ability to perform duties in
  connection with the operation of the aircraft. The certificate or
  license must have been issued or rendered valid by:

  (a) The State in which the aircraft is registered; or

  (b) The State of the Operator, provided that the State of the Operator
  and the State of Registry have entered into an agreement under Article
  83bis of the Convention on International Civil Aviation that covers
  the aircraft.

Since a captain can always fulfill the duties of a first officer, then yes, someone who normally occupies the left seat can occupy the right seat.


Do airplanes have a maximum headwind gust component when landing? An aircraft POH generally gives a maximum demonstrated crosswind component but I have not seen any indication with regards to maximum gust component.

I understand that one is supposed to increase the approach speed by half the gust component but I'm not sure if there is a limit to the gust component above which operation becomes unsafe.

Example: imagine a GA aircraft with a typical takeoff speed of 60 kts, approach speed of 70 kts. Runway is 27, wind is 270 degrees at 20 kts. What would be the maximum gust speed along the runway to operate in safe conditions?

Note: this other question is similar but focuses on crosswind component.

This is quite similar to the linked question but does have some differences. Gust components while reported differently are still part of the crosswind component. In other words a reporting like

  Rwy 03, wind 090/25 gusting 40 kts

Is another way of saying the wind is holding at 25Kts and could be as high as 40Kts at times while you are on approach. In other words you may have a 25kt cross wind on touch down or you may have a 40Kt cross wind. Generally speaking gusts occur in the direction of the wind. In some cases gusty days also come with shifty wind conditions which add for a whole new element of fun.

The maximum gust component is not quoted in the POH because there are to many inconsistent factors that may effect it. The main one being the duration of the gust. A short gust well in excess of your crosswind maximum may be recoverable while long gusts in excess of your crosswind maximum may be unsafe. Strictly speaking a loose limit would look something like

(maximum reported crosswind) + (gust wind * gust time) > cross wind component

but gust time is not reported and thats a lot of trig to crunch on the fly so no limit is typically given. I have landed a Piper Archer in conditions gusting in excess of 40Kts safely but the wind was no more than 20 degrees off the runway.

Eventually you just run out of control surface...

  Example: imagine a GA aircraft with a typical takeoff speed of 60 kts,
  approach speed of 70 kts. Runway is 27, wind is 270 degrees at 20 kts.
  What would be the maximum gust speed along the runway to operate in
  safe conditions?

In this scenario, with the wind down the runway, your maximum gust component is 50Kts. Above that and at your 70kts approach speed you would be moving backwards across the ground making it quite hard to reach the runway and potentially difficult to land the airplane if a gust were to occur at a low altitude. Most trainer aircraft like the one you describe cant do better than ~120Kts so 100Kts is about your gust limit otherwise its going to be hard to get to the runway.

NOTE: Gusty conditions (in my experience) are also generally accompanied by turbulence and in bad cases, wind shear. In many instances all of these factors combined can make an approach unsafe or un-flyable.

There is also an interesting discussion on it here.


Why can’t more older 737s be retrofitted with more newer winglets? Some background: there are four different types of winglet that have been used on various 737s over the years.

The mini winglet is used to retrofit 737-200s (one of the two 737 Original variants, the other being the 737-100).  It is basically a tiny trapezoidal tab riveted to the rear portion of the wingtip.

(Image by Julian Whitelaw at airliners.net, via meisam nemati here at AvSE.)

The blended winglet is what most people think of when they hear the word “winglet” - an attachment that looks like someone simply extended the wingtip and then bent it up to near-vertical.  It was factory-installed on all 737 Next Generation variants (the 737-600/-700/-800/-900) for the first two decades of production, and has also been retrofitted on most of the 737 Classic models (the 737-300/-400/-500) that remain in service.

(Image by Ralf Roletschek at Fahrradtechnik auf fahrradmonteur.de via Wikimedia Commons, modified by Altair78 at Wikimedia Commons.)

The split-scimitar winglet has replaced the blended winglet on the last few years of the Next Generation production line; it is shaped somewhat like the blended winglet, but a) is swooshed rearwards at its upper tip, as opposed to the simple trapezoidal tip of the blended winglet, and b) has an additional scimitar-shaped, rearward-swept fin projecting from the outside of the upbend in the winglet.

(Image by Mnts at Wikimedia Commons.)

The MAX winglet is used on the 737 MAX series (737 MAX 7/8/9/10); it, too, is a split winglet,1 but of a simpler shape than the split-scimitar winglet, looking much like someone extended the wingtip outwards, then split its upper and lower surfaces away from each other and bent them upwards and downwards, respectively, to form what looks remarkably like a glorified Airbus wingtip fence.

(Image by Aka The Beav at Flickr, via Helmy oved at Wikimedia Commons, modified by Altair78 at Wikimedia Commons.)

Now for what I’m curious about:

Although the 737-100 (the original Original) was produced in smaller numbers than the 737-200 and is no longer in service, it still hung on with various operators past the turn of the millennium; why weren’t any -100s also retrofitted with mini winglets, given that the -200 is basically just a stretched -100 with more fuel capacity?
Given that the blended winglet, despite being originally designed for the Next Generation wing, was retrofitted on the vast majority of the Classic fleet with little to no trouble, why couldn’t it also be retrofitted on the remaining Originals, given that the Classics are essentially CFM56-engined, (sometimes) stretched -200s?
Seeing as the Next-Generation-designed blended winglet was easily adopted for the earlier Classics, what keeps the split-scimitar winglet, likewise designed for the Next Generation wing, from also being retrofitted to the Classics?

(Yes, I know that this is three questions rather than one, but I felt it was better to have one question asking three closely-related questions than to clutter up the homepage with three separate but near-identical questions.)

1: As for why Boeing went to the trouble of developing two different styles of split winglet for the two in-production 737 families, that’s a different question.

Winglets change the spanwise loading on the wing. By increasing the effective span of the wing, lift is increased towards the tips of the wing. With more lift there, you get an increased bending moment on the wing.

All this means is that you can't just put winglets on or change the ones you have. You must retrofit the structure of the wing which is very expensive and requires significant downtime. It also adds a weight penalty which decreases the benefits of the winglet.

For example, on the A320, which was already wingtip fence equipped, going to the sharklets requires removing the skin and reinforcing both the tip and the center wing box near the fuselage. This takes the aircraft out of service for 15 days and involves, I've heard, in excess of $1 Million in parts alone.

On the 737, the list price for the blended winglets is $1.06 Million.

So the answer for why is because it often is not economical to do such a huge modification given the limited lifetimes on older models.


How to remove clogging in aircraft fuel nozzles? I have a dual orfice fuel nozzle (Used in CFM56 Engine/ Parker make). The secondary orifice is not opening up even at the required injection pressure. It had been not used for 3 years and I assume that it is clogged. Please suggest a method to clean this.

Hoppes No. 9 is often used for cleaning aircraft fuel injection nozzles in general. This solvent is widely available for cleaning firearms. The cleaning procedure typically involves submerging the detached nozzles in the solvent for as long as necessary, typically 10-60 minutes depending on the level of fouling.

Disclaimer: I don't have any specific maintenance data on your nozzles, so you may want be wary of the applicability of the information I offer.


What is the difference between Cruise and Service ceilings? I am looking at aircraft performance data. There are charts for "Cruise Ceiling" and "Service Ceiling."  What is the difference?  (The cruise altitude is lower.)

Ceiling refers to the density altitude the aircraft is capable of reaching under a set conditions (configutaion etc.). The main difference between service and cruise ceilings is the aircraft rate of climb at those altitudes. At service ceiling, the rate of climb is 100 fpm. From FAA pilot handbook:

  The service ceiling is the altitude at which the aircraft is unable to climb at a rate greater than 100 feet per minute (fpm).

The cruise ceiling, on the other hand, is the altitude at which the maximum climb rate is 300 fpm, though I've not seen it used much. The absolute ceiling is the maximum altitude where the aircraft can sustain level flight. From the same document:

  At the absolute ceiling, there is no excess of power and only one speed will allow steady, level flight.

The image below shows the data for A-10.

Image from forums.eagle.ru

Note that these values depend on the aircraft configuration.


Is rebooting the computer normal before/during flights? On a recent flight the captain made an announcement at the gate that they need to reboot the computer as somethings aren't working, and rebooting usually solves the problem.

After posting this on twitter I got a replying saying that most planes have 3 computers and rebooting during flight is common.

What are the reasons for reboot? For me it sounds like a software/computer that isn't tested thoroughly instead.

Today's modern airplanes have many computers, and sometimes things get out of sync and don't communicate properly.  During the preflight checks, we check each system according to the recommendations of the manufacturer, and occasionally one of the checks will not pass correctly.  Depending on the system, when a check fails, rebooting the computer may be an option to see if it fixes it (similar to rebooting your home computer).  Other times, the issue is the communication protocol between two (or more) different computers.  Often times in this case, the entire airplane must be powered down in order to reset them all at the same time.  Since these tests aren't performed in the air, the problem usually doesn't crop up there.

In the vast majority of cases it isn't a safety critical item that we are dealing with (where the software testing is more thorough), but rather a monitoring system.  The safety critical items (like flight control software, etc.) will have multiple redundant systems and often have different versions of software on at least one of the computers "just in case".

Generally the flight crew won't tell the passengers when they need to "reboot" a computer unless we have to power down the entire airplane.  That's very noticeable and will cause some people to worry so we like to give you a heads up before all of the lights go out and the air conditioning system shuts down.  That of course is up to the discretion of the flight crew though.


Why do planes employ vortex generators if the laminar flows exhibit a far lower skin friction drag than turbulent flows? I am a student and still relatively new regarding fluid flows. Our professor asked this and I can't find a source to answer this question.

Laminar flows detach from the surface earlier than turbulent flow does, which means that with everything else constant, a laminar wing stalls at a lower angle of attack, and does so more suddenly and with less warning.

The unsteady nature of a turbulent boundary layer allows it to extract more energy from the freestream (flow which is not slowed down by the wing), thus the boundary layer stays energized and slows down later. This, as you have mentioned, increases drag, so there is a tradeoff between delaying stall and increasing drag.

You'll find VGs on the wings of trainer aircraft,so that they have forgiving stall characteristics...

...STOL/bush aircraft, for lower stall speeds...

and near control surfaces of all kinds of aircraft, to retain control when the control surface itself is in high angle of attack. This also increases their effectiveness so the designers can get away with making them smaller.


Are there any helicopters with ejection seats? Are there any helicopters with ejection seats? If so, how is the clearance problem with the rotating blades solved?

Many years ago I read about the Kamov-50 helicopter family* being the first helicopters equipped with ejection seats. And Wikipedia verifies my memory:

  It is the world's first operational helicopter with a rescue ejection system, which allows the pilot to escape at all altitudes and speeds.

In the same article for the same helicopter it is mentioned that

  Before the rocket in the ejection seat deploys, the rotor blades are blown away by explosive charges in the rotor disc and the canopy is jettisoned.

Also, if you want to count experimental helicopters in, Sikorsky S-72 RSRA may be another one. Here is an interesting video, demonstrating the blades being blown away, and ejection of the pilots.

As a side note, in the article I read in an aviation magazine many years ago, it mentioned a synchronization of the seat with the blade rotation so that the seat passes between the blades. It was based on a working principle similar to the middle machine gun that some WW2 German airplanes were equipped with that fired the projectiles between the propeller blades.

Judging from the fact that I read it somewhere in the late 90s where the project was still new and developing, as PlasmaHH pointed out, that was probably more a thought rather than a fact.

* Kamov 50 and its variants/successors like Ka52


Why would Boeing be under review if crew-error is determined to be the cause? The captain on Lion air's flight 610, re-trimmed the aircraft through 21 MCAS activations. The first officer, who was given control of the aircraft, got 5 more. Although it is speculated that neither pilot executed the required stabilizer trim procedure, the report does not specifically state whether or not that procedure was actually run. Source

The captain on Ethiopia Airlines flight 302, tried engaging the auto-pilot, on the faulty side as well, and despite that being the first step for the procedure, to disengage, he tried multiple times to do otherwise, then skipped over the second step, which was to disengage auto throttles, this allowed their airspeed to build up around 390. I’m sure this didn’t help the nose down force, but ultimately, they re-engaged the system which allowed MCAS to drive the nose down further. Boeing's CEO stated these failures to follow procedures is what led to the crash. Source

The question, then, is this: if the crashes each resulted from pilot error, why is Boeing under so much scrutiny?

In aviation (and other safety critical areas) it's not enough to just look at the first cause of an accident, you have to look at why that mistake was made, and why that thing happened, as far back up the chain as you can. Then you try to address all the problems.

So, when a pilot makes a mistake, it's tempting to say "the pilots was an idiot. Case closed." However, then you've got to wonder why the company employed idiots, and how many  others are flying. That implies that there's something wrong with the pilot licencing and testing rules, and that has huge implications.

If we assume that these pilots weren't idiots, we have to ask why they made a mistake, because other good pilots could make the same mistake. Anyone who has played a computer game knows that mistakes are inevitable when you're under enough stress, just like a well built bridge will still break if it's overloaded.

You could conclude that they hadn't been trained well enough, which raises the question "Why?". Apparently the pilots did all the required training, so maybe the training was at fault. But it's not enough to just improve that piece of training, you've got to ask why again. Was it not checked? What else hasn't been checked properly? How do we know that any future training will be properly checked? Also, what incentive did the person writing the training have to miss things out, and does it apply elsewhere?

If you don't find and fix the root of a problem, it will continue to grow until it pops out somewhere else and causes another tragedy.


Is high propeller pitch inherently inefficient? For an airfoil, in general, the aerodynamic force is more or less perpendicular to the chord line.
A low-pitched propeller blade has it's chord line more perpendicular to the direction of flight, therefore, more of the force acting on it can have a greater component parallel to the direction of notion, and less perpendicular. If this force is acting forward, then this is efficient. But this may not be possible at high speeds and low rotation velocities. Thus:
A high-pitched blade has it's chord line more parallel to the direction of flight than the above. Therefore, the aerodynamic force acting on it is likely to have a greater component perpendicular and lesser component parallel to the direction of motion. This is always inefficient.
The above manifests itself in the higher-pitched propeller producing more energetic swirl than the lower pitched propeller.
So my conclusion would be that all else being equal, high-pitched propellers are inherently inefficient. Is this correct? Is this what is observed in practice?

Your conclusion is too simplistic. On aircraft, the difference between a cruise propeller and a climb propeller is rather small. However, in other applications, a low pitch propeller gives great acceleration but the engine will reach max RPM faster, which reduces top-end performance. A high pitch propeller will result in slower acceleration but higher top speeds.
This answer is also admittedly very simplistic -- there is a universe of math and engineering that goes into propeller designs. Sensenich, for example, produces more than 100 fixed-pitch propellers for aircraft, airboats and UAV's. McCauley Propellers produces thousands of constant-speed and fixed-pitch configurations for commercial, military, UAV, and general aviation applications.
If you want to get into the nitty-gritty of this, I have been told this book is one of the propeller design bibles for aircraft applications, "Aircraft Propeller Design" by Fred Weick, published in 1930.


Why did upward turning propellers in the F-82 Twin Mustang cause a loss of lift in the center wing section? In the Wikipedia article on the F-82 Twin Mustang it says:

  After a month of work North American engineers finally discovered that rotating the propellers to meet in the center on their upward turn created sufficient drag to cancel out all lift from the center wing section, one quarter of the aircraft's total wing surface area. The engines and propellers were then exchanged, with their rotation meeting on the downward turn, and the problem was fully solved"

Why did this direction of propeller turn cause this problem?

Image source: http://www.au.af.mil/au/afhra/wwwroot/photo_galleries/aaf_wwii_vol_vi/Captions/078_Experimental.htm
(This image or file is a work of a U.S. Air Force Airman or employee, taken or made as part of that person's official duties. As a work of the U.S. federal government, the image or file is in the public domain.)

I'm not sure the history provided in the Wikipedia article is entirely accurate, for the fact that both engines in an outward spinning, two-prop counter-rotating configuration are critical.  This is confirmed in the introduction of this book, though it seems to be written well after the fact and without reference to the oft-quoted story of not getting enough lift on its maiden run:

  Turning the propellers toward the fuselage meant much better control during single-engine operation.

Likewise, the republished F-82 Operator's Manual boasts, "Single-engine control characteristics of this airplane are exceptionally good."  Again (but also as expected), no mention of any trial and error in the design phase.

And the more I researched the F-82's design, I found various and unsourced information.  Many sites repeated the initial flight test story, while another said the first attempt's failure was attributed to weight.

My guess is that initial designs of the F-82 simply borrowed from the slightly earlier P-38's final configuration, where the outward-spinning counter-rotating propellers reduced downwash on the horizontal stabilizer and made for a more stable gunnery platform.

As for figuring out the lift cancellation in the F-82's center wing section, it is probably helpful to explain first what effect a single propeller has on the wing behind it.

Generally speaking, a spinning propeller creates a downstream vortex in the same direction as the propeller is rotating.  The area of wing behind the upswing of a propeller blade experiences a higher angle of attack to the resultant airflow, resulting in increased lift and drag across that section of the wing.

Image source: Civil Aviation Authority of New Zealand

Likewise, the opposite occurs on the downswing side, resulting in a net increase in lift and net decrease in drag behind the engine as a whole.  This single-engine airflow model is generally true for multiple-engine models with all propellers spinning in the same direction:

  The interaction of the slipstream with the wing produces an increase in overall lift coefficient. This net increase is the result of a local increase in local lift coefficient $C_L$ for the part of the wing located in the slipstream with up-rotation while the downrotation induces a local decrease in $C_L$. The local drag coefficient $C_d$ is locally increased in the uprotation part and decreased over the down-rotation section.  The net result is generally a reduction of drag on the wing.

So the F-82 seems to be an enigma in this regard.  Theoretically, an outward rotation should have increased lift on the center section, not reduced it.  So why didn't it?  Perhaps there was too much interaction between the two propeller wakes.  Again, consider the P-38 Lightning.  Though they both have twin counter-rotating engines and two tail booms connected by a shared horizontal stabilizer, the P-38 has a big center fuselage for the pilot sitting right between the two engines, effectively isolating the individual propeller slipstreams.

A study on the "Down Between Engines" (DBE) configuration of the Airbus A400M noted "[strong] interact[ion]" of its counter-rotating engines' slipstreams

  especially [in] the part of the wing located between both engines and results in a local decrease in lift coefficient that is not entirely offset by the local increase on those parts of the wings that experience upward motion (emphasis mine)

resulting in both decreased lift and increased drag in the wing section between the engines.

Note that DBE (i.e. inward-spinning propellers) was the fix for the F-82's lift woes; that the inter-engine wing region still experiences decreased lift and increased drag in the more preferable counter-rotating configuration underscores the notion that interaction of propeller slipstreams is not desirable.  My guess is that this switch removed just enough of the adverse lift conditions so as to enable the aircraft to get airborne.  It would be interesting to find out if anyone modified an F-82 to remove the counter-rotating aspect.

UPDATE:
At long last, I have found what appears to be the most credible source for the Wikipedia article, a 2013 issue of Warbirds magazine.

The first flight was (at least somewhat) attributed to being overweight.  A second attempt on 16 June 1945 went unexpectedly airborne for over an hour after taking off with a limited fuel load.
The interaction between the propellers created an early stall condition forward of the wing when combined with the normal upward flow, causing the lift over the center wing to predictably suffer.  According to NAA aerodynamicist Ed Horkey:

  "What was happening was that we had propellers rotating in different directions on the left and right engines.  For some reason, which I can't remember, we started with the blades moving upward in front of the center section.  What this does, particularly at high angles of attack, is to create upward flow approaching the leading edge of the center section of the wing.  You also have normal upflow ahead of the wing.  The two upflows would add together to create an early stall.  The center section represented a large portion of the wing's area.  What was happening was that we were stalling out early and just not getting enough lift. ... We changed the rotations to go down the middle and we had the problem solved."


Why are turboprop intakes placed below the propeller? On many turboprop aircraft, such as the A400M and the ATR 72, it seems that the air intake tends to be placed below the propeller (images: Wikipedia):

source

source

Given that mounting the intake in a higher location helps reduce the risk of foreign object damage from ingesting debris into the core, why are the intakes still mounted below the propeller? Are the wings and engines already mounted high enough that the intake location doesn't matter when it comes to FOD, or does another issue (maybe accessibility for ground crews?) outweigh the risk of FOD?

Inertial separators are augmented by gravity when located under the engine. It is easier to discharge ingested ice and debris downward, rather than some other direction, so that is where they go if possible. Inertia keeps solid matter low at the indicated point while air makes the turn into the engine.

C130 Nacelle


What kind of delay does the A320's fly-by-wire system add? More precisely, I am wondering if there is any kind of data on how long it takes between moving the sidestick and observing a reaction on the ailerons or elevator. I know that the physical control surfaces take time to move and, thus, I am only wondering about the actual latency added by the fly-by-wire (FBW) system.

My naive assumption is that the A320 runs a fixed control loop with precise timings. So, every $n$ milliseconds it samples the sidestick, calculates the necessary control surface deflections and then sends that via some bus which runs at $m$ Hertz. So, on average there is an expected delay of $x$ milliseconds between input and the control surfaces starting to respond to the input.

Also, is the delay noticeable by pilots?

Yes there will be a delay, but the delay caused by the control loop is really tiny. I've seen position control loops run successfully and stable at a couple of hundred Hz for simulator motion systems, and the time delay is just one iteration frame = less than 10 msec.

And as @ymb1 correctly points out, if we deflect any control surface the end position is subject to actuation forces, aerodynamic pushback forces, inertia of the surface, cable stretch and -friction, aero damping, maximum flow velocity through the servo valve etc etc. The total transfer function of stick input => surface position is such that an extra time delay of 10 msec in the total loop is not really noticeable. And control position is only the input of the aircraft response, with aircraft inertia playing a large role.

So my answer would be: a tiny time delay that can be totally disregarded.


Why was the 307’s pressurisation system removed when it was modified into the C-75, despite the fuel-burn advantage of higher-altitude flight? The Boeing 307, an airliner derivative of the B-17 heavy bomber, was the first pressurised airliner to enter series production and go into service (hence its bulbous, altogether un-B-17-like fuselage, the better to hold pressure), allowing it to cruise at higher altitudes,1 where the air is thinner and less draggy, without incapacitating its occupants.  As World War II broke out shortly after the 307 was introduced, only ten ended up being produced (out of which only nine still existed, the first off the line having already been destroyed in a crash), and eight of the nine surviving 307s were drafted into military service;2 Pan Am’s fleet of three flew military contract flights in Latin America and the Caribbean, while TWA’s five 307s were extensively modified for transatlantic VIP and freight carriage (being redesignated as the C-75).

One of the modifications applied to the ex-TWA aircraft3 was the removal of their pressurisation system, ostensibly to save weight; this restricted the C-75s to an altitude safe for unpressurised human beans (approximately 8 kilofeet or so max, as opposed to the 12 kft of the pressurised 307), with resultant drag and fuel-burn penalties in the denser, lower-altitude air (one of the other modifications was the installation of numerous additional fuel tanks to sustain the aircraft across the Atlantic).

This does not seem to make sense, given that the main advantage of cabin pressurisation is the aforementioned ability to fly at higher, less-draggy altitudes, where an aircraft cruises at a (sometimes much) higher true airspeed and burns less fuel, allowing it to carry more paying payload and not as much fuel.4  Thus, even a relatively-low-ceilinged aircraft like the 307 should still have been able to carry more payload, or carry the same amount of payload farther, when pressurised than when unpressurised, even factoring in the extra weight of the pressurisation system, due to the lesser fuel requirements of (and, thus, lighter fuel load needed for) flight at pressurisation-enabled higher altitudes - so why did the USAAF remove the 307s’ pressurisation equipment?

1: Albeit not that much higher, due to its low fuselage pressure rating and its lack of a passenger emergency oxygen system.

2: The remaining 307 (the second one off the production line) was Howard Hughes's personal airplane (he hardly ever used it, though).

3: Which would be returned to TWA at the end of the war, and converted back to commercial passenger configuration (with a number of improvements gleaned from wartime B-17 work).

4: Contrary to what those less-well-informed about aviation tend to believe, the purpose of cabin pressurisation is not to allow airliners to fly above bad weather; even modern jetliners, which can cruise at up to FL410 (~41 kft), don’t fly high enough for that.  You’d have to go up to at least about FL550-FL600 (~55-60 kft) to overtop storm clouds, something of which only one airliner ever produced was capable (and said airliner has now been out of service for over a decade and a half).

Pressurization is primarily for crew and passenger comfort. The B307 was a very early attempt at cabin pressurization and the marginal pressurization differential of only 2.4 psi did not allow much of an increase in cruising altitude.

The difference between 8,000’ and 12,000’ did not really justify the added expense and weight.

When the B307s were converted to C-75 military cargo configuration, there was even less justification for pressurization as military crew and passengers would not expect the comfort of pressurization. The C-75 could still fly at altitudes up to 23,000’ but the crew and passengers would just need to use supplemental oxygen.

The B307 had Wright R-1820 engines which were similar to B-17 engines except they were only supercharged and lacked turbochargers. Cabin pressurization was supplied by compressors mechanically driven by the engine. These compressors would rob the engines of power so removing them would allow the aircraft to fly higher but with less comfort.


What physical hardware does all the avionics on the B737NG run on? Some processor from the 1980s? Or some new and improved chip? So, I am a tech guy and likes to fly (in sims) so, I was wondering what physical hardware powers all the 737NG's avionics and related systems (PFD, MFD, EICAS, GPWS etc). Is there an answer to this?

This wouldn't fit as a comment but doesn't get to the processor level of your question. The 737NG/Max has a large number of avionics systems as shown by Airframer. Most of these have processors, some have multiple. In which of these are you interested?


How is stability ensured with bulk-loaded cargo? According to this question especially smaller airlines often don't use ULD in the cargo hold. But wouldn't that be dangerous for the stability of the aircraft? Won't the cargo move and unbalance the aircraft?

You can simply tie down the cargo to prevent its movement. Most planes have nets that subdivide the bulk cargo compartment, so as to prevent movement of the baggage. For example, the following image shows the cargo hold of a B737 with dividers.

Image from airport-data.com Photographer:  Nick Taylor Photography

The ground operations manual of a carrier gives the following procedure for securing loads in A320:

  1.1.4.2.1. Load Securing A321 (bulk loaded)

  The divider and door net in each compartment must be closed at any time. Additional tie-down is normally not required except for individual items of load which by their nature, shape or density may constitute a hazard. They must be restrained, which can be achieved by filling the cargo hold or net
  section volumetrically, or by tie-down.

  ... Packages weighing more than 150kg shall be restraint or individually tied-down. Single packages should be tied-down. Tie-down of loads to aircraft structure is achieved by straps or nets connected to the tie-
  down points located on the cargo hold floor.


Has this "straight through" airport design ever been contemplated? Here's an airport, with some planes on the various runways/taxiways and some using the terminal ..

(airliners in green)

Airport/terminal design is really tricky.

There are many spatial problems which are unsolved, resulting in the incredibly inconvenient (for humans/planes) current airport paradigm.

Has an "over" terminal ever been contemplated?

terminal indicated in blue

So, the (whole) airport terminal building is simply up on stilts.

(Obviously, the clearance is high enough for the tallest existing and forseeable airliners.)

At a stroke this utterly solves two problems:

Airliners never again need to reverse.  They just land, drive around anywhere they want, stop under the terminal.  Passengers get on and off.  The airliner drives off again, drives to a runway (perhaps again passing under the terminal - no issue) and flies off.

No craft would ever reverse.

This would be a complete paradigm shift, slashing handling times, costs,  personnel, machinery.

The fundamental spatial problem of airports is solved at a stroke: You can land or take off anywhere, and get to and from, any gate.  (Sure, you have to avoid the few pillars holding up the terminal.)
From the human crowd movement viewpoint, the need for the hell of "separate terminals" is slashed or even eliminated. At worst you'd just have two or three "ranks" of these (rather like giant rows of car parking in a parking lot).

Again this would spectacularly slash the size and infrastructure costs of airport concept design. Often the whole concept of ridiculous people movers, monorails etc would be eliminated or slashed.

Note that the car freeway system would just connect to the end(s) of the terminal building. (Indeed, as usual.)

(But indeed, perhaps the blue structure could indeed include "the road itself". So the blue object represents both car driving lanes and the terminal structure, shopping, etc.)

On that: looking ahead, it's possible there could finally be a drive to your actual gate concept (if check-in / baggage handling / border functions could be rationalized), which would be an total paradigm shift in air travel.

The only technical difficulty is AeroTech or whoever would have to make a jetway that comes "down"...

Note that it's obviously trivial to make huge civil structures "elevated",

(Indeed a civil structure for foot traffic, shops, etc is (I believe) a lighter load than a super-freeway or bridge,

)

Ideally the supporting piers would be infrequent, allowing for totally free and easy movement of airliners, but even if the columns were between each "bay" the major paradigm shifts are achieved.

This airport paradigm would seem to be staggeringly more efficient than the current paradigm.  No reversing ever!  Free driving anywhere for all craft over the entire area.  Wasted space slashed.

Has this ever been contemplated?

Yes. An architect named Miklos Deri seems to have contemplated something similar: Drive Through Airport – Inversion of Procedures. The design is based around smaller aircraft (which reduces some of the height and size problems) and stacks aircraft together in lines to address the space efficiency issues.

Nobody has built such an airport (it's a concept, not created for any particular project), and there are practical problems associated with moving aircraft multiple times during a turnaround.


How precise are altimeters? How precise and how accurate are typical altimeters these days? That is, if two planes are flying (side-by-side, very near each other), and both altimeters show the same Flight Level, what is the maximum vertical difference in actual altitude you would expect to see between them? (I.e. is it on the order of a few meters, a few tens of meters, etc?) Is there any legal standard for precision/accuracy of the altimeter that must be met (let's say in the US)?

While Accuracy and Precision are closely related, they're not the same thing.
In the context of your question both are important:

Accuracy

In the US altimeters are broadly required to comply with FAR 43 Appendix E limits for accuracy (technically this is regulatory for IFR flight, but TSO'd "sensitive altimeters" will generally meet these requirements - at least when they left the factory - and that's what you're going to find in most aircraft).

The accuracy requirements provided in FAR 43 are based on mechanical altimeters, and are subject to a number of requirements (friction error limits, hysterisis limits, etc.), but the test tables there are a good guide to the accuracy of "typical" altimeters.

Also note that the accuracy of sensitive altimeters depends on the pressure setting the pilot enters (an incorrect pressure setting could mean substantial error). In "the flight levels" everyone uses the same standard altimeter setting (29.92 inches of mercury / 1013 millibars) to avoid this issue (similarly the ATC transponder transmits an altitude assuming a pressure of 29.92 inches of mercury - ATC's computers correct that value based on local altimeter settings as required).

  Depending on the design digital altimeters may also be immune to some of the accuracy-affecting issues in an analog/mechanical altimeter - for example a solid-state digital altimeter should not suffer from "friction error" as there's no rotating/moving shaft involved.

Precision

Analog altimeters are "reasonably precise" -- take this example:

The tick marks are every 20 feet. One could easily eyeball 20 foot increments, and even 10 foot increments (with the pointer directly between two ticks). Similarly the pressure settings are marked every 0.02 inches of mercury (or 1mb), allowing for a fairly precise setting of the value.

Digital altimeters are extremely precise (often offering one-foot precision in their display, and 0.01 inches of mercury for the pressure setting).

  I believe the precision requirements for altimeter markings are specified in the relevant TSOs, but I don't have those documents handy.

In both cases the precision of the markings can exceed the accuracy of the device: At 10,000 feet the FARs allow an 80 foot tolerance, and whether your measurement is presented as a needle position (roughly to the nearest 10 feet) or a number (to the nearest foot) you could still be 79 feet higher or lower than you "think" you are and be "within tolerance".

So, combining all of this together - If we have two aircraft flying along indicating the same altitude (say 10,000 feet) and both altimeters are perfect there will be zero feet of vertical separation between them (the static ports of each aircraft will be at the exact same altitude).
On the other hand if both aircraft are at the outer edge of tolerance (one +80ft, one -80ft) there will be 160 feet of vertical separation between them (discounting visual "slop" in the analog altimeter and assuming the pointer is perfectly on the line and the pressure setting is perfectly dialed in).

(In either case the answer would be "WAY TOO CLOSE!")


What is a Cessna 172's maximum altitude? What is the maximum altitude at which a Cessna 172 can fly?

The absolute ceiling, or the maximum height an aircraft can fly to, is usually not published; manufacturers usually use service ceiling as the benchmark.  Service ceiling is the maximum altitude which the aircraft can attain flying in air at Standard Temperature and Pressure (29.92" Hg and 15° C at MSL) and still be able to climb at a rate of at least 100 feet/min.

Depending on the model 172, this is between 13,000-15,000 feet ASL.  A Cessna 172SP has a published service ceiling of 14,000 feet ASL.

Again whether an airplane can attain these altitudes on any given day is dependent on the condition of the air.  The aircraft may be able to ascend higher than this on days when the air is very cold and there is a high local barometric pressure or lower on days when the air is warm and/or at a low barometric pressure in the region.


What is the horizontal and vertical fields of view for a VASI and a PAPI approach system as you come in to land? As you come in for landing, at what angles should the VASI or PAPI lights be visible on the approach in the Horizontal and vertical planes. {E.G. 30 degrees horizontal and 20 degrees vertical}

This information is for the US, taken from FAA Order 6850 - Visual Guidance Lighting Systems.

VASI (2 bar VASI for illustration, but other configurations exist)

a. Layout

b. Aiming

  The two-bar VASI system shall have the upwind bar aimed at three
  degrees and the downwind bar aimed one-half degree lower at 2.5
  degrees; however, where necessary to provide obstacle clearance, the
  downwind bar may be aimed at 2.75 degrees.

c. Glide (vertical) path angles

d. Horizontal field of view: 10°

  Determine the minimum aiming angle by plotting all objects within an
  area  established by two lines each diverging at 10 degrees from
  runway centerline extend to 4 nautical miles, and starting at the
  runway centerline at the number one bar.  A line drawn from the runway
  centerline at the number one bar at an angle that will clear the
  highest object by not less than one degree establishes the minimum
  aiming angle.  This line and a horizontal line perpendicular to the
  runway centerline at the number one bar (downwind) establishes an
  obstruction - clear plane 10 degrees (6 degrees in special
  applications where 10 degrees is not obtainable) on both sides on the
  extended runway centerline.  The obstruction clear plane starts at the
  number one bar and extends outward from the threshold for four
  nautical miles.  In addition, the minimum aiming angle shall be within
  the glide path limitations of paragraph 4. When less than 10 degrees
  are used for determining the obstruction-clear plane, issue a NOTAM
  and request publication in the Airport/Facilities Directory.

PAPI

a. Colocation with ILS glideslope

  When siting PAPI on a runway with an established electronic glide
  slope, the PAPI visual approach path should coincide, as much as
  possible, with the one produced electronically.  To accomplish this,
  the PAPI is placed at the same distance from the threshold as the
  virtual source of the electronic glide slope within a tolerance of ±30
  feet (±10 m).  The PAPI is aimed at the same angle as the electronic
  glide slope.

b. Layout

c. Aiming

  The visual glide path angle is the center of the on-course zone, and
  shall normally be 3 degrees when measured from the horizontal.

d. Glide (vertical) path angles

e. Horizontal field of view: 10°

  The PAPI obstacle clearance surface is established to provide the
  pilot with a minimum clearance over obstacles during approach.  The
  PAPI must be positioned and aimed so that no obstacles penetrate this
  surface.  The surface begins 300 feet (90 m) in front of the PAPI
  system (closer to the threshold) and proceeds outward into the
  approach zone at an angle 1  degree less than the aiming angle of the
  third LHA from the runway.  For a 3 degrees glide path and 20 minutes
  separation between LHAs, the third LHA from the runway would be aimed
  at 2 degrees, 50 minutes elevation.  The surface extends 10 degrees on
  either side of the runway centerline extended, and extends 4 statute
  miles from its point of origin.  The surface is shown graphically in
  figure 5-4.  If a site survey determines that there is an obstacle,
  which penetrates the obstacle clearance surface and cannot be removed,
  then the glide path angle must be changed or the PAPI system moved
  further from the threshold.  By moving or re-aiming the PAPI, the PAPI
  obstacle clearance surface is repositioned so it will not be
  penetrated by an obstacle.


Formula for calculating turn indicator (deviation from coordination) being more a professional programmer and less prof pilot, I am building a small
Android EFIS and I'm able to calculate quaternions, yaw, pitch, roll, TAS, acceleration values etc. I am stuck calculating the deviation from a coordinated curve indicator, usually the small ball in the turn indicator. Could someone please point me to a formula? I'm finding only formulas about calculating the necessary bank angle or radius from speed.
Thanks in advance and best regards
hk

I am stuck calculating the deviation from a coordinated curve indicator, usually the small ball in the turn indicator.

The "small ball in the turn indicator" (which pilots use to determine what they need to do with the rudder pedals in order to stay in coordinated flight) is just an inclinometer. It's literally just a heavy ball in a curved tube filled with oil.

If you want to replicate the way that the inclinometer works, then look at the lateral acceleration and vertical acceleration and use the arctan2 function to determine the apparent tilt.

By "lateral acceleration," I mean the proper acceleration in the direction of the aircraft's lateral axis, since that's what matters here and that's what you're able to sense. Likewise, by "vertical acceleration," I mean proper acceleration in the direction of the vertical axis.

The oil in the tube functions as a low-pass filter with a time constant of maybe about a second.

All that said, I don't think there's actually any reason to use the vertical acceleration. You can just take the lateral acceleration, low-pass filter it, and display the result.


What are the minimum requirements for passing medical for PPL? Where can I find the UK requirements to pass the medical exam for a PPL?

Basically I would like to check my health and all the requirements with any doctor - I have health insurance plan, and when I know that I meet all the criteria I go back to the aviation doctor for checking and issuing the certificate.

The UK aviation medical standards are available online, so I assume you could go to any doctor with that information for a checkup.


Why does an elevator deflection (step) result in a nonzero short-term-steady-state response in pitch rate? I know from the simplified Short Period linearized equations (state space with Angle of Attack ($\alpha$) and Pitch Rate ($q$) as states) that the steady-state value of the pitch rate response to elevator deflection (step) is non-zero

I can also relate that to the spring-mass system and understand that the concise derivatives $M_{alpha}$ and $M_q$ work as spring and a damper, respectively.

I think (would appreciate confirmation) that the term short-term steady state should mean that the phugoid motion is going to take over eventually, being the dominant mode.

However, I would like to understand that physically, because $\alpha$ and $q$ seem to be both on steady state (impossible unless flight path changes). Also, I would expect $q$ to return to zero at some point. Does it mean this equation is oversimplified in this regard?

The equation is certainly oversimplified, hence the name 'short period': it only models short period pitch oscillations. In this model, there are only two states: pitch and pitch rate. Any other effect is thus neglected.

A step response in an elevator would in the short term definitely lead to a nonzero pitch rate, as the model predicts. In reality, the pitch rate will usually decay to zero, because the airspeed drops to zero (there may also be altitude density effects). Imagine what would happen if you would pull on the stick in an aircraft, and crank the throttle to make sure the speed does not decay. You would do a vertical loop, exactly like the $q$ term predicts.

Of course, then you run in to an extra complication: your model is linearized. The concept of large angles, let alone a loop, does not exist in the linearized world. Think of a linearized pendulum: that works fine for small angles, but as far as the model is concerned, a 360° angle means just a very large deflection, and not the reality of returning to zero deflection after a full loop. The model is then only valid for the small amplitude oscillations it is intended to predict.

A phugoid mode exchanges speed for altitude and vice versa. Neither of these states are represented in  your short period linearized model. This behaviour is impossible to see in your model. You might as well ask why you don't see the effect of the Moon position in your model; it's simply not in the state space!

A long story short then: simplified models like these are great tools, but cannot and should not be used outside of what they are supposed to do. If you want to estimate instantaneous pitch rate from elevator deflection then this is the model for you. If you want to check for trim conditions, large attitude deviations, or even roll behaviour, then you need a better model.


Where is the first Boeing 747-8 ever built stored at today? I have tried many different ways to try to find where it is stored, but have had no luck.

The first 747-8 built was not a protype, but a production unit. The first delivered aircraft in both the versions (747-8i Intercontinenal and 747-8F Freighter) are in service, the Passenger version with Lufthansa and the cargo version with Cargolux.

The first 747-8 customer was Cargolux for the 747-8F version.

The first 747 ever built, 747 RA001 'City of Everett' is in Museum of Flight in Seattle.