Can the Airbus A380 safely fly with two engines out on the same wing? I know twin-engine jets can safely take off and fly with one engine out (i.e. all thrust only on one side).  Of course quad-engine jets are also able to fly safely in case one engine is out.

But is the A380 designed to also fly safely in case two engines on the same wing are out?

Would the rudder be able to compensate the thrust imbalance?

I'm referring to an engine-out situation in flight, not during take off.  I would assume that the problems aggravate during take off since the rudder has less airflow to counter the imbalance.

Note: Two engines out on the same wing is not a very likely scenario, but it's not impossible.  For example, on Qantas 32, damage from the number 2 engine explosion impacted fuel and control lines of engine 1.

Proving that a multi-engine aircraft is flyable with one or more engines inoperative is part of the certification programme. The Airbus A380 is certified under FAR (Federal Aviation Regulations, USA) part 25 and JAR (Joint Aviation Regulations, Europe) part 25. One of the requirements of FAR/JAR 25 is that the directional control can be maintained when two critical engines (that means under the same wing) fail.

The following is selectively extracted from FAR part 25 (JAR is similar, there are some differences). Emphasis added by me.

  Controllability and Maneuverability

  §25.143   General.

  ....

  b) It must be possible to make a smooth transition from one flight
  condition to any other flight condition without exceptional piloting
  skill, alertness, or strength, and without danger of exceeding the
  airplane limit-load factor under any probable operating conditions,
  including—

  The sudden failure of the critical engine;
  For airplanes with three or more engines, the sudden failure of
  the second critical engine when the airplane is in the en route,
  approach, or landing configuration and is trimmed with the critical
  engine inoperative;

  ...

$ $

  §25.147   Directional and lateral control.

  a)  Directional control; general. It must be possible, with the wings
  level, to yaw into the operative engine and to safely make a
  reasonably sudden change in heading of up to 15 degrees in the
  direction of the critical inoperative engine. This must be shown at
  1.3 VSR1 for heading changes up to 15 degrees (except that the heading change at which the rudder pedal force is 150 pounds need not be
  exceeded), and with—

  The critical engine inoperative and its propeller in the minimum
  drag position;
  The power required for level flight at 1.3 VSR1, but not more than
  maximum continuous power;
  The most unfavorable center of gravity;
  Landing gear retracted;
  Flaps in the approach position; and
  Maximum landing weight.

  b) Directional control; airplanes with four or more engines.
  Airplanes with four or more engines must meet the requirements of
  paragraph (a) of this section except that—

  The two critical engines must be inoperative with their propellers
  (if applicable) in the minimum drag position;
  [Reserved]
  The flaps must be in the most favorable climb position.

  c) ...

  ...

  d) ...

  e) Lateral control; airplanes with four or more engines. Airplanes
  with four or more engines must be able to make 20° banked turns, with
  and against the inoperative engines, from steady flight at a speed
  equal to 1.3 VSR1, with maximum continuous power, and with the
  airplane in the configuration prescribed by paragraph (b) of this
  section.

  f) ...

$ $

  §25.149   Minimum control speed.

  ...

  g) For airplanes with three or more engines, VMCL-2, the minimum control speed during approach and landing with one critical engine
  inoperative, is the calibrated airspeed at which, when a second
  critical engine is suddenly made inoperative, it is possible to
  maintain control of the airplane with both engines still inoperative,
  and maintain straight flight with an angle of bank of not more than 5
  degrees. VMCL-2 must be established with—

  ...

  ...

$ $

  §25.161   Trim.

  ...

  e) Airplanes with four or more engines. Each airplane with four or
  more engines must also maintain trim in rectilinear flight with the
  most unfavorable center of gravity and at the climb speed,
  configuration, and power required by §25.123(a) for the purpose of
  establishing the en route flight paths with two engines inoperative.


How does p-factor affect jet engines? How does p-factor affect jet engines, if at all? If it doesn't, why?

Since P Factor is primarily variation of AOA of a thrust producing blade when its rotation axis is not aligned with the air stream, you should see some P factor in a turbofan engine, to the extent that the inlet flow strikes the front of the fan from an angle.
I say "some" because the cowl inlet is going to align the inflow to be more perpendicular to the fan face than if the fan was in the freestream (where it becomes just a propeller with lots of extra blades), but it's hard to say how much. In any case, the flow is probably not quite 90 degrees to the fan face at higher body AOA, so it's safe to say there is probably a little bit, but since P factor is not accounted for in turbofan installations, what P Factor exists is not enough to matter (maybe if you stuck a turbofan on the nose of a Cessna 172, you might notice some P Factor effect).
On a turbojet there will be no P Factor at all since the thrust is coming form flow accelerating out the burner can, hot section and tailpipe, not from lift producing propeller/fan blades at the front.


What is "Carson cruise" and can I determine it myself? Mike Y. just posted an answer containing the following:

  With today’s fuel costs, the Carson cruise angle of attack, which represents the most efficient way to fly fast with the least increase in fuel consumption, is of significant interest.

Being an outstanding human, Mike also added a reference to the document that it came from:

  "Low-Cost Accurate Angle-of-Attack System", Borja Martos and David F. Rogers, Journal of Aircraft 2018 55:2, 660-665

But for those of us who don't have access to that journal or can't be bothered to go find it, what is Carson cruise?  What exactly are the benefits and how is it determined?  Finally, can I determine it myself for my airplane without an AOA indicator?

From this highly reliable web page http://www.eaa1000.av.org

source

The maximum range airspeed may be the most economical way to fly, but it's quite slow... Hence, we may consider the speed for which the ratio fuel flow/knot of airspeed reaches a minimum. That speed, marked by the point where a line from the origin is tangent to the thrust required curve, is called 'Carson's speed'.

source

In s/l flight, drag = thrust. Hence, the lower graph shows the 'thrust required' curve. There is a fixed ratio (1,316) between the maximum endurance speed, maximum range speed and Carson's speed, so that:

Maximum range speed = 1,316 x Maximum endurance speed

Carson's speed = 1,316 x Maximum range speed

That's a simple way to find Carson's speed for your plane. You can get the value of the Max. range speed from the POH


Would a refueling stop be cheaper than finding lodging for all the passengers? This is somewhat late followup question to my original question. Basic story was flight got delayed by almost 24 hours because first, technical reasons, secondly temperatures were very high.

As far as I can see airlines had to accommodate all the passengers somehow, either in terms of hotel rooms and meal or on other flights. There are costs involved in every action.

If they had taken off Abu Dhabi on less fuel and then refilled somewhere else would it have been cheaper? Would such refueling stop be allowed where passengers are not allowed to get off? Just fill her up and continue.

Without knowing any technical/legal/contractual details, my simplistic view is, you land at some airport around 6 hours away, fill up your tanks - might take around 1 hour and continue rest of your 14 hours flight. Your crew could be in time limits as they must be provisioning for full length of the flight plus possible diversions, moreover almost working as planned and you don't have to pay for hundreds of rooms and/or rebooking/re-routing costs of the passengers.

I don't know what aircraft you flew on over 3 years ago. The current options from Etihad are Boeing 777 and Airbus A380. The route doesn't seem to be nonstop anymore, aboard any airline.

Along the shortest distance -- great circle route -- there aren't many options for an A380 to land and find support equipment mid-route (shown below). Refueling can occur with the passengers on board, there is a procedure to be followed, which includes alerting the fire service at the airport to be ready.

If we take only the cost of fuel into account, and a 777-200 as a baseline, actually stopping every ~6 hours is more economical. (This doesn't take into account landing fees, accrued cycles*, passenger convenience, etc.) Not just for your flight, but any long-haul flight; the fuel burned to carry the fuel for long distances becomes considerable after a certain range.

Another consideration is which freedoms of the air, say Etihad/UAE, has with say Russia. The second freedom specifically is:

  The right to refuel or carry out maintenance in a foreign country without embarking or disembarking passengers or cargo.

While they are called freedoms, not all country-pairs mutually allow all the freedoms. The specifics are usually not public and are contained in the air service agreement between any 2 nations.

In all, it might have saved fuel even if we included a detour off the great circle. It certainly would have increased the maintenance cost. And it would have been an operational nightmare to arrange compared to a hotel and meal.

* Pressurizing and depressurizing twice, applying the takeoff thrust twice, applying the brakes twice, etc., all increase the maintenance cost.

Re: comment about London or Sweden already being destinations:

Already having flights to Sweden, UK, or Italy, doesn't mean the 2nd freedom is guaranteed. Note that "not many options" $\neq$ none. The middle half of the flight is either over tundra or the Arctic Ocean. An airport like London Heathrow won't welcome an on the spot request for a slot; busy airports in Europe are running near max capacity, so unless it's an emergency, on the spot slots really aren't realizable. Three daily slots to land at London Heathrow (as a destination) already cost Etihad £46.2 million back in 2013.

(gcmap.com)


Is there a specific relation between sideslip angle and yaw rate OR how sideslip angle depend on yaw rate? Since, yaw rate is controlled via rudder and sideslip angle can also be controlled by rudder, I want to know if there is any dependence of sideslip angle on yaw rate in fly by wire aircraft as I am working on the SAS design of a/c lateral/directional dynamics?

Since, yaw rate is controlled via rudder

No, yaw is controller by many things, and rudder is there mainly to compensate for them to keep the plane flying nose-forward.

The usual way to make a control system is that you control (with a PID controller):

Move elevator to target vertical (plane coordinates) acceleration.
Move ailerons to target roll rate.
Move rudder to target zero lateral (plane coordinates) acceleration. Lateral acceleration is measured by the ball in traditional cockpit.

Note that each axis is basically independent and rudder is only used to eliminate side-slip, which is proportional to the lateral acceleration unless you have asymmetric thrust or drag—and I have not heard of a FBW that would automatically correct for those conditions.

When flying with auto-pilot, the auto-pilot sets vertical speed target from altitude error and vertical acceleration from vertical speed error for the pitch and it sets roll from heading error and roll rate from roll error for roll axis. Autopilots don't set lateral acceleation targets at all—that is always zero.

That is why auto-land has much lower cross-wind limit—autopilot does not handle rudder, so it can't de-crab.


Were any composites used on the A340 series? Were any composites used in the construction of the A340? If so, which ones?

Yes!!

The slide shows that composite are used in rear pressure bulkhead and keel beam in addition to other parts.

Source: https://company.airbus.com/dam/assets/airbusgroup/int/en/investor-relations/documents/2008/presentations/GIF2008/gif2008_workshop_composites_hellard.pdf

You can read more:
http://www.lcc.mw.tum.de/fileadmin/w00bkg/www/PDF/Symposium/LCC_Symposium_Havar.pdf


What is the coordinate system used in an Aircraft? I know that from nose to tail the aircraft is divided by STA (station). At least for Boeing and Airbus this is the case but I am not familiar with other manufacturers.  However, there is also two other dimensions that give every location of an aircraft an identifiable point location. It's obviously not x,y,z.  What is this coordinate system and approximately what does one unit along each access mean in terms of inches or centimeters?

Stations (STA), generally Fuselage Stations (FS), designate location along the length of the aircraft, increasing from nose to tail. Generally station 0 is somewhere in front of the airplane. One reason for this is that if the airplane grows longer, there still won't be any negative stations, and sections of the plane keep their station numbers. FS 100 can also be placed at an important location (looks like the firewall below).

Water lines (WL) designate location in the height of the aircraft, from ground up. You can probably guess by now that these are holdovers from the marine industry. Like the stations, water line 0 is generally a bit below the aircraft, possibly to have WL 100 at the center of the fuselage, or WL 0 could be at the bottom of the fuselage.

Butt lines (BL) designate location left/right on the aircraft, generally centered in the middle.

While the main system is based on the whole airplane, different parts (like the wings) have their own system of STA, WL, and BL.

The stations are generally in whatever unit is being used for the aircraft. In the US, this is inches.

See this link for an example and further explanation.

For a picture similar to the one below but with more detail, see here.


Why are delta wings not commonly used in commercial airliners? Delta wings are most commonly seen on military aircraft. I understand that delta wings are ideal for supersonic flight, and provide high maneuverability.

Having high maneuverability and speed are good characteristics for commercial airliners. But, most of the commercial airliners do not use delta wings (Except for Concorde and Tu-144).

Are there any specific reasons for this?
Are there any regulations that prevent using delta wings in commercial airliners?
Does it cost more in maintenance?

There's a thread on airliners.net about the same topic.

  One of the most prominent and useful aspects of the delta wing is the fact that it increases an airplane's critical Mach number (the speed at which some air passing over any part of an aircraft - usually the wing - reaches Mach 1.0).

By increasing the critical Mach number, it makes supersonic flight much more feasible on delta wings than on conventional wings.

Delta wings are particularly happy even at a high AoA (angle of attack). This is because of the reaction lift generated by the wing, and the vortex created at the leading edge of the wing. Hence, airplanes with delta wings have a much higher tolerance to stalling. This makes it easier to attain higher vertical speed, also known as "rate of climb" safely.

But on the flip side:

  One of the downsides of deltas is that they create more drag, because of the bigger surface area, but they do have a bigger internal volume (for fuel, for example) which could lead to a smaller fuselage, or more space in the fuselage for other things. Another disadvantage is higher landing and take off speeds.

The last point is particularly important. The Concorde's already small number of usable airports was reduced because the runways needed to be long enough to accommodate for the 160 KIAS landing speeds (for comparison, typical turbofan-based airliners lands at 130-140 KIAS).

As a result of the higher takeoff/landing speeds, the  cost of maintenance of vital components such as the brakes and tyres is higher. Moreover, the high landing speeds can even lead to greater stress on the airframe. So, indirectly delta wings do contribute to a higher cost of maintenance.

Other disadvantages of delta wings I can think of is the lack of a flaps/slats system. Delta wings don't require flaps as they can simply reduce their airspeed by increasing the AoA, but the lack of flaps reduces their subsonic efficiency, and increases the takeoff distance and takeoff velocity. For instance, the Concorde's delta wings produced very little lift during takeoff prior to rotation, necessitating high takeoff speeds of nearly 220 knots!

As an answer to your second question, I could not find anything about regulations relating to the use of delta wings on airliners.

All in all, delta wings work well for niche purposes such as supersonic aircraft, but their disadvantages and poor operating economics make them impractical for subsonic flight.

However, as you can see here, the now-scrapped Boeing Sonic Cruiser attempted to make the best of both worlds by combining a conventional wing with a delta wing, to form a wide, "cranked-delta" wing design.


What are those 5-character alphanumeric waypoints? On the following High IFR chart from skyvector.com I notice along the meridian at 30' intervals there is a green symbol the same shape as an RNAV waypoint. Next to them are 5-digit identifiers, but it is alphanumeric rather than the usual 5-letter fix name. What are these symbols and what are they used for? Can they be entered into the FMS?

That's part of the High Altitude Redesign (HAR) program.

As far as the available information, it's now available for flights above FL390, and only in certain ARTCC's in the U.S.

The waypoint naming convention is as follows:

There are plans to have this system implemented worldwide.

The pilot will no longer be limited by airways based on ground navaids. I.e. they'll make their own airways based on those waypoints. It's called Non-Restrictive Routing (NRR).

This way they can plan to fly around the weather, yet remain close to the great circle route. The waypoints are also easy to relay and eliminates confusion when re-routing, as the identifiers are based on their location.

Comparison to the ground-based navaids is shown below:

More information:

FAA AC 90-99, and NRS Description.


Why is it not a good idea to climb higher than a level at which you would have a cruise weight-climb speed less than 500 ft/min? In relation with What is weight-climb speed?, why is it not a good idea to climb higher than a level at which you would have a cruise weight-climb speed less than 500 ft/min?

Because ATC expects you to climb with at least 500 fpm when crossing to the next level. Otherwise you will block intermediate levels for an extended period of time.

If you are climbing 2000ft to the next available level for your direction of travel, that will take at most 4 minutes at 500fpm or more. This means the flight level you are crossing will need to be free of traffic from the opposite direction for four minutes to ensure legal separation is maintained. If you climb slower, ATC can't properly accomodate your climb.


What is the meaning of "Docile pitch behavior" in the context of Pugachev's Cobra? In another answer on SE Aviation it is mentioned that the ability to perform  Pugachev's Cobra  in a non-thrust-vectored aircraft needs "docile pitch behavior of the airframe up to approx. 110° angle of attack"

Can an F-16 perform Pugachev's Cobra?

What does "docile pitch behavior" mean in this context? Does it mean that when the control surfaces are returned to neutral the aircraft has a natural pitch down moment that returns it to its horizontal position?

If so, does the pilot actually have control over how long he can hold the vertical position in Pugachev's Cobra? What control surface enables him to exercise this authority in the post-stall regime when I'd have assumed the
control surfaces lack such authority?

Or is the holding time of the vertical attitude in Pugachev's Cobra beyond the pilot's control and only related to some sort of "natural period" of this maneuver in that aircraft linked to how long it takes the restoring moment to return the aircraft to its nose down attitude?

"Docile" means no big pitching moments over the angle of attack range. Since the control power with fully stalled control surfaces is rather small, the airframe must not create strong pitching moments, but keep a small pitch-down moment so the airplane returns eventually to normal angles of attack which allow regular flight.

The maneuver is dynamic, so the pilot needs to build up a specific pitch rate. Inertia will keep the aircraft pitching up, and the small pitch-down moment (and aerodynamic damping) will reduce the pitch rate so that the pitch-up stops when the nose is pointing straight up. What follows is the consequence of the small pitch-down moment: The aircraft slowly returns to controlled flight, and once control authority is re-established, the pilot stops the pitching motion.

By selecting the initial pitch rate, the pilot can control the maximum angle of attack, but once he leaves the flight regime of attached flow, he has to wait until inertia and aerodynamics do their part.


What are the advantages of towbarless pushback tugs? After watching the video below I wondered the reason for towbarless tugs. Lifting the front wheel of a 575 ton aircraft seems like it requires an awful lot of expensive and complex hydraulics when the goal is just to move the aircraft around on the ramp. What benefit is there to doing this over just using a towbar?

Advantages of towbarless pushback.

You don't have to store different tow-bars for different a/c types. This is especially good for ground handlers who handle different airlines with different aircraft types. Also handy for ad-hoc one-off flights.
Removing need for tow-bars means one thing less to check daily, weekly, monthly, annually.
Towing with a tow-bar requires a man in the cockpit to ride the brakes (brake-rider). This also means the aircraft may require power (either from APU or from the towing vehicle) to supply hydraulic pressure. An aircraft towed with a towbar is dependent on the aircraft brakes if a sudden stop is required as the towbar will not be able to take the weight of the aircraft on its own on decelerations.
A normal tug (which uses towbar) needs to be very heavy to enable it to tow an aircraft. Without the weight, the tug will just spin its wheels. The tow--barless tug will have the weight of the aircraft on its wheels so it doesn't need to be as heavy. Just solidly built to support the engine and associated hydraulic stuff. This weight also makes the brakes more effective.
A towbarless can tow the aircraft faster as control is way better. In some busy airports they only allow you to move the aircraft across busy taxiways or during peak hours if it is either taxied under its own power or towed by a towbarless tug. Towing using tow-bar is not allowed as it is very slow.
The tow-bar has 'shear pins' which limit the force which can be applied through the nosewheel towbar points. This reduces the towing speed. They can and do break once in a while. While a spare is usually carried in somewhere on the 'bar it still stakes a few minutes to get in installed. Not fun in the middle of a pushback and may result in a blocked taxiway.
A lot of injuries are caused by towbars. Either they jacknife when something fails or a lapse in procedure caused the nosewheel steering to be activated before the bar has been disconnected. The danger here is the nosewheel is self-centering in almost all aircraft and will swing the bar if the nosewheel is not centered. Scary to see and can take your legs out.

I'm sure there are more aspects but these are the major points. Hope this clarifies a bit.


How to calculate DDM in a ILS dual frequency system? I'm working in a project where I have to process ILS baseband signals. I have no problem obtaining DDM value from single frequency signals, nor the individual components of a dual frequency signal (DDM from clearance signal and DDM from course signal).

Now I need to merge those two individual DDM values (in dual frequency systems) to obtain an unified DDM value, similar to what Rohde equipment does (see figure below).

I've been told that I can obtain this
$\sum DDM$ value either:

As a function of individual DDM values: $\sum DDM = f(DDM_{clr}, DDM_{crs})$. But I'm not able to find the actual function $f(·)$.
Mixing clearance and course signals at baseband, so both components add up resulting in only one ILS-alike signal, and then calculate DDM from this new signal.

Are any of these methods valid? If not, how can I obtain $\sum DDM$ in a dual frequency system?

I don't think you're supposed to just average the two together somehow by themselves. The two signals are designed to be used in different zones and in practice have very different signal strengths at any point on approach.

From EVALUATION OF INSTRUMENT LANDING SYSTEM DDM CALIBRATION ACCURACIES [sic] by Dennis M. McCollum [PDF]:

  The course pattern contains sidelobes that, without the clearance pattern, could cause an aircraft to select a wrong (sidelobe) course. Within the region of the true course (i.e., + 10 degrees about the designed procedural course), the signal strength due to the course array is greater than that from the clearance array. In the regions where the sidelobes exist, the clearance signal predominates. Airborne ILS receivers are designed to respond to the greater of the two signals.

This has to do with the background in processing these signals through analog means. Both the clearance and the course signals are within the passband of the receiver.  Since this is an AM receiver and DDM is more or less a ratio of 90 Hz to 150 Hz, if the clearance or course is much stronger than the other, the stronger one prevails.

So in practice measuring the course signal DDM inside the clearance region doesn't reflect what's done in practice, and I'd guess it probably isn't even a guaranteed behavior of the transmitter.  Do anything else instead in this region, like perhaps measuring just the field strength of the course signal.

But what about when slightly off the centerline?  Then you'd get a mix of course and clearance signals.  You could assume the ideal condition where a capture effect completely eliminates the weaker frequency.  Another approximation presented in the admittedly confusing paper "Current Issues in Demanding ILS Ground and Flight Measurement Environments" by Gerhard Greving and L. Nelson Sponheimer is

\[ DDM_{total} = \frac {DDM_{clearance}* CR^x + DDM_{course} } {1+CR^x} \]

Although I can't quite make out what precisely CR and X is here.

Depending on the application, it may be more realistic to model the amount of distortion expected or permissible due to clearance band interference, especially due to multipath reflections of that signal.  However, given your comments that probably won't help your specific application.


Do fuel tank locations on the wing cause a potential for weight imbalance? I have a question about fuel tanks in an older plane like a P-36C Hawk and a P-26B-35. The fuel tanks in the P-26 are placed in the wings and in the P-36C Hawk its under the pilot and behind him. For the P-26 or any plane with its fuel tanks on the wings would it cause more weight on one side if the fuel is used more then the other?

YES

If the port and starboard tanks drain at different rates in a gravity fed system (due to a blocked valve or a clog in a fuel line), then eventually a differential in remaining fuel will drive a lateral imbalance in the plane's weight.  Another (and better) way of saying this is that the center of gravity will no longer be aligned with the aircraft centerline.

I can't comment on how this situation might emerge in the P-26 specifically, as I could not find a detailed enough description of its fuel system.

It is worth noting that the same is true with fore/aft center of gravity drift in aircraft where the tanks are located entirely in the fuselage (or under it).  This often has a more dangerous effect if not properly managed (and as a result the Cg drifts too far aft).

It is also worth noting that lateral asymmetry can also be created by non-simultaneous release of stores (like the 200 lb of bombs that the P-26 carried).  However, in the case of the Peashooter, the bombs were carried very close to the centerline, so the asymmetry would be minimal (as compared to that on a P-51 for instance).


Is there a way to get the original data used to create performance charts? I want to create a flight planning software for GA. Just as a prototype for the time being. The performance data and info about the planes should be stored in various databases.

When I go through flight manuals like the CESSNA 150, I get a lot of tables which I could convert into a database, but where do I get data which created these charts e.g. the range profiles?

Is this data available somewhere? Or are there databases available to be used?

As others have stated, the data certainly exists but the manufacturers are under no obligation to provide anything more than what's in the manual.

That said, the performance equations are straightforward and with some effort you could build the tables.  What you are proposing exists for most air transport aircraft for their Flight Management Systems.

As a reference, you might check out Getting to Grips with Aircraft Performance.  It's written by Airbus but provides a very good explanation including much of the math.


How does the static pressure system keep the ambient pressure? I understand that the compressed airflow around the aircraft effectively lowers the air pressure, and that's the reason why the alternate air switch affects your altimeter — it uses the slightly lower pressure in the cabin as reference — or why fuel is sucked out of the tanks if you forget your fuel caps.

This begs the question, what in the design of the static pressure system makes it more closely match the ambient pressure? Why is it unaffected by the effects affecting the cabin pressure?

Test. Calibrate. Repeat.
There is a new method developed by NASA based on GPS tracking. And here's the same thing but as a video (also NASA).

New Method for Pitot-Static Calibration
A precise, time- and cost-effective method based on global positioning system technology using output error optimization (...) enables near real-time monitoring of error in airspeed measurements, which can be used to alert pilots when airspeed instruments are inaccurate or failing.

Webpage also mentions the current methods that include the trailing cone (the one you see on modern airliners while they are being tested), tower fly-by's, and pacer airplanes.

Trailing cones were first developed and tested in the 1950s and 1960s as a simple means of calibrating the static pressure error of an aircraft's pitot-static system.

So apart from it being time consuming, it should be relatively easy. If it says I'm doing 100 KIAS but I know I'm only doing 96, just calibrate the 100 to be 96. Probably needs a screwdriver and a certified mechanic. But doable. Same for the altimeter and so on. Altimeter would be easy to calibrate on ground. Put in the current QNH, check the reading against the actual (from an airport diagram) and take out your screwdriver (not literally).
The pitot-static system doesn't know the right pressure, it just needs calibration. Which also answers why the cabin atmosphere won't be as precise, it wasn't calibrated for that.


Can the ILS glideslope be intercepted from above? Is it recommended to intercept the ILS glideslope from above? Are there any special considerations like unstable approaches or GPWS warnings?

Assuming that you're asking about intercepting the glideslope from above rather than the localizer, the answer is that it is definitely NOT recommended.

There are at least two significant problems if you do this. First, is the fact that at high vertical angles there can be false glideslopes. Looking at the diagram below, if you're coming at the glideslope from above and the needle starts to move, are you intercepting the real 3° glideslope or the false 9 ° glideslope?

The diagram above is the best I could find, but it has one error. It shows a "Flag no signal" at 6°. Actually what you'll have there is another false glideslope, but one which has reverse sensing.

You can, of course, do a reasonability check if you know your distance from the runway, or you could compare your vertical speed if following it with what you know your vertical speed should be, or on the 6° line a command to go up when you know you should go down would be a hint, but it's unwise in my opinion to put yourself in a situation that increases your work load that much during an instrument approach.

You asked about GPWS warnings. None of the aircraft I flew that had GPWS would sound a GPWS warning just from receiving a false glide slope. However, all of them would have sounded the GPWS if you followed a false glideslope and were close to the ground because doing so would give you an excessive sink rate.

A second problem with approaching from above is that the vertical angular width of the glideslope as measured from needle full down to needle full up is very small. As I remember only about one degree. Whatever it is, though, the point is that if you descend into the glideslope with a vertical speed of only a few hundred feet per minute greater than the glideslope itself is descending, the needle is going to go from full down to full up very quickly. Without working it out, I'd guess that if you're inside the outer marker you might well miss its movement if your instrument scan had you away from the needle more than a second.

In the late 1970s a friend of mine was flying a charter in a Cessna 310. He unwisely decided to intercept the glideslope at Salem, Oregon from above. The aircraft hit the ground in the vicinity of the outer marker, killing all aboard. During the accident investigation, the NTSB found from radar data that he had a descent rate greater than 1,000 fpm going through the glideslope.


Why isn't the APU a standard diesel generator? Why isn't the APU a standard diesel generator? Jet-A and diesel are interchangeable (with the addition of a lubricant). So why not use a cheaper piston engine as opposed to a jet turbine?

V.S.

Weight.

Piston APUs for trucks are designed for frugal and quiet operation. This one produces 5.2 kW electrical power and weighs 375 lbs.

The APU for the A320 and B737 is a noisy screaming unit that produces 90 kW electrical power, 445 shaft kW. It weighs 375 lbs as well.

The main difference is in the weight of the engine itself. The turbine engine of the A320/B737 APU produces 445 shaft kW and weighs 145 kg. This diesel engine also produces 450 shaft kW and weighs 1542 kg.

Power-to-weight ratios:

The turbine produces 3.1 kW/kg engine weight.
The diesel engine produces 0.3 kW/kg engine weight, it is a standard industrial machine and not optimised for weight.
The Rolls Royce Merlin was very much optimised for weight, it used aviation fuel (gasoline) and produced 1.58 kW/kg


How much unusable fuel is in the A320? I was wondering how much unusable fuel is in each tank or maybe in total for an A320. I searched through the FCOM and could not find answers.

To add to this question is the fuel shown on the fuel SD page all usable fuel? Or is it unusable + usable fuel?

Please show the references to the answers.

You can find the amount of unusable fuel in the EASA Type Certificate Data Sheet. Here is an example table for the A320-200 from that document:

In general, the cockpit displays will always show the amount of usable fuel only. This is mandated e.g. by EASA CS 25.1337:

  (b) Fuel quantity indicator. There must be means to indicate to the flight-crew members, the
  quantity, in litres, (gallons), or equivalent units, of usable fuel in each tank during flight. In
  addition –

  Each fuel quantity indicator must be calibrated to read ‘zero’ during level flight when the
  quantity of fuel remaining in the tank is equal to the unusable fuel supply determined
  under CS 25.959;
  Tanks with interconnected outlets and airspaces may be treated as one tank and need
  not have separate indicators; and
  Each exposed sight gauge, used as a fuel quantity indicator, must be protected against
  damage.

(EASA Easy Access Rules for Large Aeroplanes)


How do you request a "pop up" IFR clearance? If you are flying VFR into deterioration weather conditions and would like to continue the flight under IFR, or you would like to climb to a better altitude but are restricted by a cloud layer, what is the easiest way to go about requesting and instrument clearance.

Assuming you are not already talking to ATC:

You: Houston Center, Cessna 1234X, request*

Center: Cessna 34X, Houston Center, go ahead

You: Cessna 34X is a 172/U, 35 North IDU, 4500 ft, request IFR to Austin

Center: Cessna 34X ident, fly heading 270

Center: Cessna 34X is cleared to AUS via radar vectors, climb and maintain 6000 ft, squawk 5634

This is essentially the same as making a request for VFR flight following services with the exceptions:

Your aircraft needs to be legal for IFR.
You need to be rated and current for IFR.
You will receive a clearance.

If you already have flight following and need IFR, the controller has you on his scope and already has a strip for you, and you just need to ask him to make you IFR.  In that case he will just put you on an IFR altitude and give you a clearance limit and route.

*As noted in the comments, you may decorate this call with something identifying the nature of your request


Can two opposite localizers to a single runway be on simultaneously? In case of a single runway with two separate ILS systems (one in each end, facing localizers) operating on two different frequencies, can both localizers be on and transmitting at the same time, or must one be switched off if the other is on?

Please refer to ICAO recommendations if possible.

ICAO Annex 10 Volume 1 has the answer:

  3.1.2.7 At those locations where two separate ILS facilities serve opposite ends of a single runway, an interlock shall ensure that only the localizer serving the approach direction in use shall radiate, except where the localizer in operational use is Facility Performance Category I — ILS and no operationally harmful interference results.

In plain language: only one localizer can be on at a time, unless the ILS in use is only CAT I and it has been confirmed that no harmful interference will happen.


Could anyone (non-military) in the present day "exceed mach 1 over a regulated area"? There is a question What would happen if I exceeded mach 1 over a regulated area?

I wonder, could anyone in civil aviation (not military) actually carry out this violation? I know there are privately owned "warbirds", but it looks like the majority of them are really old. The "never exceed" velocity is defined well below Mach 1 for majority of these planes.

Are there any cases known where someone non-military owns an aircraft that could exceed the speed of sound and then land at least mostly in one piece?

Let's limit the scope of the question to USA and Europe.

Yes, In the U.S.A there are many privately owned ex-military jet fighters that are capable of exceeding Mach 1. It is quite conceivable that a privately owned fighter jet has exceeded Mach 1 but not reported it or been caught by the authorities.

You can search the FAA civil registry by type:  FAA REGISTRY
Make / Model Inquiry

Here is just a partial list of privately owned supersonic jet fighters in the United States:

Lockheed F-104
McDonnell F-4
NORTHROP F-5
MIKOYAN-GUREVICH MIG-21
MIKOYAN-GUREVICH MIG-23
MIKOYAN-GUREVICH MIG-29
Sukhoi SU-27

These two aircraft are owned by the Collings Foundation in Florida and are available for airshow demonstrations.


How to calculate camber for a balsa glider? I am making a rubber powered balsa glider for a competition. I want to make the glider fly as long as possible. I've heard that a high camber means slower speed means more airtime. I want to make sure that the camber isn't too high, and I need to know what type of camber I should use. Also, I need to know if the horizontal and vertical stabilizer size is fine.

The dimensions of my glider include: 32cm fuselage, 7.8cm wing chord, 28cm wingspan, 5.5cm horizontal stabilizer chord, 11.5cm horizontal stabilizer span, 5cm vertical stabilizer height, and 5.5cm vertical stab width.

Welcome to Aviation SE, Samarth!

Flight duration depends not only on camber, but also launch speed, launch angle, mass, propeller diameter, propeller pitch, rubber quality, winding tightness, etc.  I assume that by "rubber powered glider" you mean rubber band powered for the first part of the flight, gliding thereafter.

Even if you somehow find optimal values for all the other parameters except for camber, it's still simplest -- and most trustworthy! -- to build several interchangeable wings for the glider, varying camber, and directly measuring flight duration with a stopwatch.


Aerodynamic advantage of blunt noses and WHY (subsonic) Compared to a sharper conical nose, a blunter conical nose is aerodynamically superior in subsonic flight. My question is why. I've read a lot about this and allegedly:

A blunter nose accepts a wider range of angles of attack without sideways airflow across it, which would cause additional drag that is avoided.

It would appear that airliners have blunter noses because they fly at a wider range of AOAs and gliders have sharper noses because they fly at a narrower range of AOAs. However...

At optimal AOA, the form and skin friction drag is also lower. I don't understand why!

According to this, a blunt nose is aerodynamically superior even when flying at the optimal AOA, but then why do gliders, which obviously value aerodynamics above all, always have relatively sharp noses? According to some descriptions, a hemispherical cap is the optimal shape for decreasing form drag and skin friction drag, but I haven’t found any description of why the first is true and the second is often motivated with hemispherical caps having the lowest surface area to volume ratio, but a conical cap of equal height has a lower surface area relative to the radius! In other words, if we were to design a new airplane, why wouldn't we sharpen the nose closer to a point instead of blunting it closer to a hemisphere, only considering aerodynamics?

Edit: yes, this is a unique question because I’m asking why and specifically why form and skin friction drag of a blunter nose is smaller than for a sharper nose, challenging intuition.

Edit 2: basically, why does No 2 in the following have less drag than No 5?

According to the image (I’ve yet to read the source, I’m on it!) there’s less pressure drag, but according to an answer by aeroalias to a related question, there’s less friction drag:
Why/when is the blunt nose better?
Which one is it? Both? Which one is most relevant? Or are both irrelevant in reality? Peter Kämpf’s answer makes no mention of either alternative, only arguing about the AOA and stagnation point.

Nose shape on fuselages

Fuselage noses of airliners hold the antenna of the weather radar, and a blunter nose will put the least amount of material in the way of the electromagnetic waves. On a three-dimensional body the shape of the nose is of less importance and sometimes used to express a brand image. In subsonic flow, sharp changes in the contour must be avoided: Here, air would have to change direction abruptly which requires strong pressure gradients. Therefore, curvature along any flow path should change gradually. This favors an elliptic body shape. If you want to have no jump even in the second derivative, use a lemniscatic function.

A hemispherical nose would have constant curvature followed by no curvature along the cylindrical portion of the fuselage. The sudden change in curvature at the transition between the rounded nose and the cylinder would require a sudden jump in pressure which causes more drag than the gradual reduction of curvature (causing a gradual increase of pressure) of an elliptic nose.

  Edit 2: basically, why does No 2 in the following have less drag than No 5?

Because the sharp corner at the base of the cone in No. 5 will cause a suction peak which in turn will cause a jump in boundary layer thickness. The hemispherical nose of No. 2 is better, but best is the elliptical nose in No. 1.

A blunter ellipse will have lower overall surface but a steeper change in pressure - here, as so often, the optimum is a compromise which needs to include the Mach and Reynolds numbers as well as structural considerations. Since the optimum is flat, there is considerable space for individual solutions.

  What if No 5 had a smoother “ridge” (cone to cylinder transition) but was still pointy? Say, if No 1 ended in a point?

The pointy tip would not "fix" the stagnation point to that point - rather, it would still move with the angle of attack, but the part of the flow negotiating this tip would exhibit a strong suction peak and "smooth out" the contour by means of a local separation. This would place more strain on the boundary layer and increase its thickness downstream. Any hope for laminar flow over this nose past such a tip would be destroyed by this, of course.

The smoother ridge would be better, so the elliptical nose with a pointy tip would improve things like No. 4 does over No. 5. On No. 4 the sharp ridge will most likely produce a separation on the ridge which will make the whole body look larger to the oncoming flow. The bisection of the drag coefficient in No. 4 indicates to me that the separation is much smaller. Note that this is all only valid in subsonic flow! In supersonic flow No. 4 would look best.

Nose shape on wings

On wings, however, the nose shape is extremely important - witness the jumps im performance that sometimes were achieved by optimizing the nose shape. To get to the bottom of the wing nose shape, we need to talk about boundary layers, Mach effects and much more, so be prepared for a long answer.

Glider airfoils can afford to use small nose radii because they fly at low Mach numbers. Airliners, on the other hand, need to keep local Mach numbers low which favors a larger nose radius. If you are happy with this answer, better stop reading now.

At the Symposium Transsonicum of the International Union of Theoretical and Applied Mechanics (IUTAM) in 1964, E.V. Laitone presented something like a magic number for transsonic flow:
$$Ma = \sqrt{\frac{1}{\gamma - 1}} = 1.581$$
with $\gamma$ the ratio of specific heats of a fluid. Once the local suction peak at the nose of an airfoil reaches such a speed, lift stops growing any further. With stall speeds of around 120 knots, the early jet airliners routinely hit this limit using the 6-series NACA airfoils of that time.

Below you see the pressure distribution on a flapped airfoil near maximum lift (picture source). The specifics of the airfoil don't matter much; important is the suction peak at the nose which is made possible by a large angle of attack and a flap with a ventilated gap between wing and flap.

In such a case, a small nose radius creates a very strong but narrow suction peak because a very strong pressure gradient is needed to force the flow around the tight nose contour. If the nose is blunter, the suction peak can spread out lengthwise and becomes less peaky, simply because the trajectory change of the flow around the nose is becoming more gradual. Since there is a direct relationship between local suction and local Mach number, those suction peaks must be spread out and flatter in order to allow higher angles of attack. That is the main reason why supercritical airfoils have blunter noses - they will tolerate higher lift coefficients made possible with powerful flaps.

Initial glider airfoils used rather blunt noses and high camber. With composite technology the smoothness of wings could be improved and early numerical codes helped to shape the pressure distribution such that a fairly wide angle of attack range can be covered without incurring sharp suction peaks on either surface. The Eppler code of Richard Eppler was the first such tool and by prescribing pressure levels over sections of the airfoil at specified angles of attack it made laminar airfoils easy to design on the low-power computers of that time. The result are contours with a small nose radius which would produce a sharp suction peak once the specified angle of attack range was exceeded. By suppressing the suction peak over that specified angle of attack range, the laminar bucket could be maximized in a way that a blunter nose would not allow.

Now for that promised excursion into boundary layer theory: A laminar boundary layer is stabilized by a positive speed gradient (accelerating flow) but a negative speed gradient will trip the transition to turbulent flow rather quickly, especially at higher Reynolds numbers. The back side of a suction peak has just such a negative gradient; therefore suction peaks must be avoided in order to keep the flow laminar over much of the wing chord of gliders. In addition, using camber flaps allows to shift the laminar bucket up and down on the lift coefficient scale, so the combination of a small nose radius and a camber flap will allow laminar flow over a large lift coefficient range.

Airliners cannot do the same: The Reynolds number on their wings is so large that turbulent transition happens near the nose regardless of the local pressure distribution. Therefore, they are not concerned with laminar buckets and will happily tolerate the suction peaks that come with higher angles of attack. In cruise, when managing the pressure distribution is essential to reduce shocks, the angle of attack range is extremely narrow, so the blunt nose produces no disadvantage. However, this transsonic pressure distribution demands very little camber in the forward part of the airfoil, so in order to tolerate high angles of attack during approach and landing, they need a blunt nose. The larger nose radius also helps to integrate leading edge devices like slats and Krüger flaps, which is another advantage at low speed.


What is the difference in VTOL technology between British Aerospace Harrier II and F-35B? F-35B safer and looks more stable in terms of VTOL capability.

What is the basic difference between Harrier II and F-35B's VTOL technology?

the F35 has a ton of electronics that the Harrier could only dream about, but I will focus on propulsion differences.

The biggest is that the F35B uses a separate lift fan in front of the engine, oriented to provide vertical lift that balances the vectored thrust from the rear of the engine. The lift fan is disconnected when flight speed is obtained and covers closed. Roll control at low speed comes from bleed air from the front fan in the engine.

The Harrier uses a much larger version of a bleed system from a single engine to provide both forward located vertical lift/thrust and a similar total diversion system for rearward lift/thrust through four rotating nozzles. All nozzles are working all the time and rotated to a lift/thrust position. A separate bleed system provides roll and pitch control at low speed.


What use was there for Boeing to include customer numbers in their aircraft model names, and why did they end this practice? Inspired by this question (thanks Hot Network Questions!).
The tl;dr is that Boeing included a unique numeric model number designation, referring to the customer that had purchased a particular plane, in the actual model number of the plane. For example, a 747-8 for Lufthansa would be reflected as "747-830", where 30 is Lufthansa's unique customer number.
What is the utility of this "customer model number"? I assume it was used downstream by e.g. fitters and painters to determine what specific customisations should be applied to the airframe, but it seems that since 2016 this practice has been phased out.
Does anyone happen to know the history behind these numbers, i.e. how and why they came into being and why Boeing chose to stop using them after literally decades?

It's an easy way to know what configuration options "this" aircraft has, long after it has left the factory.
Let's say that airline 1 ordered a bunch of 737's with one cockpit jumpseat & two aircraft batteries and the 3-hour cargo fire suppression option -- stuff you might expect for overwater flights.  They got, hypothetically, the 737-7E1 (just making that up.... a 737-700 with customer code E1).  Then airline 2 ordered a bunch of 737-700's with two cockpit jumpseats & a single aircraft battery and the 1-hour fire suppression system... a more short-haul sort of fleet.  And let's say that their code was the 737-7E2.  Years later, all of those aircraft from both airline 1 and airline 2, and perhaps others, have been sold to airline 3.  Airline 3 has some 737-7E1's and some 737-7E2's.  And if you're conversant with the codes in that fleet (as mechanics & others at Airline 3 would be), then you can see on a list that "this" aircraft is a -7E2 and so you know that it's the configuration with both jumpseats, one battery, and etc.
Those customer codes don't deal with things like the passenger cabin seating configuration; those can be swapped out pretty easily.  But stuff like batteries & cargo fire bottles and plumbing for an extra O2 mask in the cockpit doesn't tend to change, so those sorts of things get captured in the customer codes.
As to why they abandoned them, my only guess is that they may have decided to standardize all their product by removing customizations -- a 787 is a 787 is a 787, and (perhaps) they don't let customers configure them in ways that, down the road, may make the aircraft less desirable for somebody looking to buy or lease a fleet that doesn't have lots of variants between jets.
There are stories from guys who flew freighters where ever aircraft on the ramp was different in some way -- even though they were all the same type.  Newer INS, older INS, different flight directors, sometimes even different types of engines, etc etc.  For a freight operation looking for any aircraft they can get cheap, that lack of standardization may be no big deal. But for other operators, not having to deal with the headaches of dissimilar aircraft in the fleet may be an incentive to choose "this" type over "that" one.


How does an increase in climb rate affect climb gradient? I'm learning to take (and pass I hope ;)) my PPL exam and I've just found one question, that I'm not 100% sure about.
How is climb gradient going to change (for small angles) while rate of climb increases 2 times? (TAS = constant)

a) stay the same
b) increase 2 times
c) increase 4 times
d) decrease 2 times

Can I calculate climb gradient as (rate of climb/TAS)? If yes, the correct answer is B.
Guys, could you please confirm?

For small angles, climb gradient can indeed be approximated using ROC/TAS. Easiest if you draw a sketch for things like this!


Is it common for a cold reciprocating engine to stop at idle? The other day I started the engine of a PA28 and put it at 1200 RPM as per the checklist but the sound didn't seem right - so I tried each magneto and the power dropped a lot on one of them. I then pulled the throttle to idle and the engine stopped. It was on for maybe 30 seconds so it was still cold.

I restarted the engine, went through the standard checklist and everything was fine the second time (I ran the magneto and idle checks once the engine was warm).

Is this behaviour expected from a cold engine (large power drop on single magneto and engine stop at idle)?

For reference, the aircraft is a PA28-180, I believe the engine is a LYCOMING O-360-A4A with ca. 500 hours. The outside temperature was 15C.

What you experienced is common, but it would serve you well to better understand your engine.

There are quite a few engine options for the PA28, so without more specifics I cannot give you specific guidance, but I will address this in a general form.

Spark plugs foul from several different sources.  The leading sources are from excessive unburnt fuel, and from oil.  The unburnt fuel can cause carbon and fuel based buildup on the plugs, or it can come in the form of lead fouling, where lead deposits bridge the spark gap or insulator in the spark plug. Oil fouling tends to happen in higher time engines, but can happen in new engines during their break-in period.  Normally, running at cruise power keeps the plugs clean by scavenging these deposits off.

Again, engines vary, and manufacturer's recommendations vary, but in general, I run with the engine lean during taxi. This reduces buildups and will clear some fouling.  Because the engine running is impacted by engine temperature, I generally do not do a run-up until after taxi, when the engine is warmer.

If an engine is running rough, or the mag check fails, or the engine repeatedly stalls, I will do a clearing run-up, where the engine is run at a higher power setting, typically 1800 to 2000RPM and leaned until it will just continue running. This creates higher temperatures, and tends to burn off oil or other fouling.  This is done for anywhere from 1 minute to about 3 minutes, and then the mixture is enrichened and a normal run-up and mag check are done.

Magnetos do not foul, rather spark plugs foul, and typically the R mag will run the top plugs on one side of the engine, and the L mag will run the top plugs on the other side. Bottom plugs tend to fould more, especially if there is oil fouling. But the point is that whether the L or R mag fails, or runs rough, doesn't tell you exactly what spark plugs it is. A cylinder temp gauge or other instrumentation may, but usually it is determined by disassembling and inspecting the spark plugs.

So to get to the OP question: Large drops at runup are a sign of either a condition which may resolve itself, or conditions which are causing the engine to run less effectively. If they can't be respoved the plane needs to be seen by a mechanic, or someone more familiar with it.  If the mag is rough, and the engine has just been started, then it is premature to perform a normal mag check.  Let the engine ward up more, during taxi, and then check the mags.  The engine should not die at idle.  If that happens repeatedly after the engine is warmer (not even up to operating temps), then there is a problem, and the engine should be checked out by a mechanic or someone with more experience.

In addition to not clue as to what engine you are dealing with, we also don't know the local meteorological conditions. If you are having problems at 0F it is more typical than having problems at 35F.  But in general, engines should not stall at idle even when cold at 40F temperatures.

A final comment. I have trained many CFIs, and I spend quite a bit of time on engines and their operation and work-arounds.  Much of this is not in FAA publications, or even training publications. I find that many CFIs are not a knowledgeable about engine operations and related methods. Being engine smart will help you in your flying, and it will reduce your risk.

Addendum:

As a general rule:  Lycoming: R Mag top left plugs, cyl 2, 4 and bottom right plugs, cyl 1, 3.
Continentals: R Mag fires the top plugs, L Mag fires the bottom. The impulse is the R Mag, as the top plugs tend to get less oil fouling.


What are the consequences of touching down at too low of a speed? The consequences of a high speed touchdown / high vertical speed are very much obvious. However, what will the consequences be of a low speed touchdown? How severe are they (if any)?

This is certainly not a comprehensive answer but can get you started. I’m assuming fixed wing non-VTOL aircraft.

A low (forward) speed touchdown, all other things being as close to nominal as possible in that scenario, translates in all likelihood to a steeper than normal landing and can be associated with loss of sufficient lift or even stall and can result in a hard landing or crash. If trying to compensate for loss of lift, the same scenario will lead to a higher pitch than normal and thus add (or replace the above with) a tail strike risk.

A low vertical speed touchdown, all other things being as close to nominal as possible in that scenario, means a shallower trajectory than normal and can be associated with a long flare or a deep landing and as such poses risks around stopping performance (e.g. if the touchdown is so soft that the automatic deceleration systems don’t even realise the aircraft is on ground) or available stopping distance (if all the runway is used up trying to stick the aircraft down, none is left for rollout). Exiting the runway down the far end at substantial speed again is a recipe for significant damage and danger for life.

A touchdown with both (too) low forward and (too) low vertical speed doesn’t really exist, as that would mean a touchdown roughly along the nominal path but with lower speed, and since approach speed is usually calculated to be somewhere around minimum safe speed (i.e. stall speed plus some margin), slowing down significantly means stall, at which point the scenario changes to low forward but high vertical speed at touchdown.


Launch control equivalent while take-off What is the proper term used when the plane is stopped and the engine is ramped up and after a certain thrust is reached, it suddenly starts off with a jerk.

The term I'm coming across the most is "static take-off" (even used by Airbus).

Its advantage is that it's using the least distance since you don't "waste" the distance it takes the engines to spool-up (in a jet anyway) to take-off power.

Not to be confused with "non-rolling take-off", which just means the aircraft was holding on the runway before the take-off run started. A non-rolling take-off is not a "static take-off" if the pilot didn't step on the brakes until the engines spooled up.


What is the flight that returned to base after dumping fuel due to toilet issues? From this post I got to know one of the flight has emergency landed due to LAV tank full and Lavatory was In-operative.

Can someone please share the flight details that made emergency landing due to Lav in-operative or Lav is full? This will help in my study (white paper publish).

I'm going to be a bit lazy and only point you in the right direction (sorry):
The Aviation Herald website is a simple, yet excellent source of aviation "occurences". There is a search tool at the top of the page, here is a link to a sample search I did using word lavatory:
Avherald.com search with word lavatory
...and below is a sample report picked from the results of the aforementioned search.

Incident: Jet2.com B738 near Cork on Aug 4th 2019, lavatory issues
By Simon Hradecky, created Monday, Aug 5th 2019 21:46Z, last updated Monday, Aug 5th 2019 21:46Z
A Jet2.com Boeing 737-800, registration G-JZBC performing flight LS-167 from Glasgow,SC (UK) to Lanzarote,CI (Spain), was enroute at FL360 about 90nm south of Cork (Ireland) when the aircraft turned around due to problems with the lavatories and diverted to Manchester,EN (UK) for a safe landing on runway 23R about one hour later. A passenger reported toilets were leaking water flooding the aircraft. The airline reported the aircraft diverted to Manchester to permit maintenance to service the toilets. A replacement Boeing 737-800 registration G-GDFV reached Lanzarote with a delay of 3.5 hours.

Searching the avherald.com database with relevant words will give you a plethora of cases for your white paper.


How can pilots sleep or rest merely 1 hour after takeoff? I screenshot this YouTube video. Because the A380 just attained FL250, the flight obviously commenced less than an hour ago.  At 14:28 the relief captain (Robert Juelicher) leaves the cockpit. At 14:37, ATC clears the flight to climb to FL320.

Doubtless, relief pilots are legally mandated to have well and sufficiently rested and slept before their flight.

So how can Juelicher  be expected to sleep or rest 1 hour just into the flight?
Or did the Pilot-in-Command (Juergens Raps exhibited below  in the screenshot) mis-speak? Is Juelicher expected to merely fritter away time (e.g. read a book) until his shift, not necessarily to sleep?

The military has been teaching people how to fall asleep fast for a long time and its quite doable. Keep in mind this pilot did not start working one hour ago, if he is relived after 1 hour he is likely at the end of his duty shift and may have been flying for the previous 7 hours and awake for some more time than that. Pilots hop time zones frequently so just because its daytime outside the aircraft does not mean that pilot started off his duty time in that time zone or even on that continent that day.

The relief pilots may have woken up just before the flight and had a nice hour to relax before "going to work".

The crew rest area is pretty nicely isolated and free of light and distractions making it an ideal place to sleep.


Can the thrust of a turbofan be increased by increasing bypass ratio? I read that the CFM56 is based upon the F101.  Id on't know how true that is, but it generated the following question:  If a larger fan is fitted to a low bypass turbofan, would dry thrust be increased?  Or is it the same amount of thrust produced with less fuel consumption?  That is what I was told on another message board but it doesn't seem to make sense.  Assuming that the proper size turbine is fitted, the same turbine airflow should be turned into more thrust with a larger fan.  Which is correct?

GE used the core of the F101 engine as the core of the CFM56 (with some modifications, presumably). The core of a turbofan consists of the high pressure compressor, the combustion chamber and the high pressure turbine. In this case, we talk about a 9-stage HPC and a single-stage HPT. Everything else was newly developed for the high-bypass turbofan.

Unlike just putting a larger fan on the low-pressure shaft, this is possible becasue the high-pressure components and the low-pressure components are relatively independent in a turbofan, i.e. the core of the engine is not affected too much by the size of the fan, the bypass ratio or other changes in the low-pressure area.

The new low pressure turbine (4-stage instead of 2-stage) can operate at a lower rpm and transmit more of the exhaust gas energy from the engine core to the larger fan (the exhaust gas coming out of the new LPT is probably cooler and slower than that coming out of the F101 LPT). The new low-pressure compressor increases the overall pressure ratio and therefore increases the amount of energy that goes to the low pressure compressor and fan, thereby improving the overall efficiency. This is probably possible because the non-supersonic operational envelope of the CFM56 means that intake air total temperatures will be much lower so that the same EGT margin can be achieved with a higher pressure ratio. In other words, becasue the incoming air is cooler, it can be compressed further without damaging the high-pressure turbine because it is too hot.

The static thrust of the CFM56 is 19500 lbf, which is up about 15 % from the F101's 17000 lbf without the afterburner. While a small part of this is probably an aerodynamic loss due to the unused afterburner chamber, most of it has to come from the better use of the energy offered by the engine core exhaust in the larger fan. The difference might even be higher becasue the maximum thrust of the F101 might be adjusted for a much lower mean time between shop visits. I'm not sure about that though.

This shows that the main benefit of the high-bypass low pressure section of the engine is a significantly increased fuel efficiency, an increase in the static thrust and (much) lower noise levels. The F101, on the other hand, can operate at higher speeds, produce more thrust at high speeds, is lighter and has a much smaller cross-section.

(Edited per comment by fooot, thanks for finding the source.)


Is an aircraft capable of flying at 900 km/h at 8000 metres above ground? There was a question in my maths textbook and I wondered if it was actually possible.

P.S. 8,000 metres is about 26,000 feet.

It is possible, but it is quite fast.

900 km/h is 486 knots.

Assuming no wind and a standard atmosphere, 486 knots ground speed is 486 knots true airspeed (TAS), which is equivalent to a calibrated airspeed (CAS) of 334 knots at 8000 meters.

At that altitude, the maximum operating speed (Vmo) is usually the limiting factor; if you are a kilometre higher then the maximum operating mach number (Mmo) will typically become limiting.

334 is above the Vmo of some airliners (they will need a little bit of tailwind to achieve 900 km/h over ground at 8000 meters), but many are capable of achieving it without wind.

Some examples of Vmo:

Boeing 737: 340 kts (350 kts for the Original's)
Airbus A320 family: 350 kts
Boeing 747: 365 kts
Boeing 777: 330 kts
Airbus A330: 330 kts
Airbus A340: 330 kts


Does it exist? Environmentally-friendly/safe fluorescent liquid to make sea crashes easier to find? Has this been tried in aviation history?

To make airplane sea crashes easier to find, can we use a colored liquid that is lighter than water, environmentally-friendly and safe, with fluorescent properties so that if an airplane crashes into the ocean, this liquid would leak out and float to the surface. During the day, the color is noticeable by satellite and by night it glows.

Does something like this exist? Can it be easily implemented? Has it already been attempted?

Thanks to those who commented and provided answer-worthy comments. Thanks!

What is already in existance today:

@Andy commented:

  For spacecraft splashdowns (certainly Gemini, maybe others) they used
  a green marker dye, but I don't know how long it lasted. That would be
  a good thing to start with as they must have tested how long the
  traces were visible when choosing it.

@RogerLipscombe commented:

  prestodye.com/sea-dye-markers says "30 to 40 minutes", depending on
  sea conditions. Apparently there's a standard: MIL-S-17980D.

And for the challenges:

@abelenky commented:

  Rough approximation: How many gallons of this stuff would you need for
  it to not dissipate in minutes? How much would that weigh? How often
  to airplanes crash in the sea? The cost and complexity just doesn't
  even come close to the benefit.


How does an inexperienced pilot handle G-Forces? After reading this question and the linked article, I had some more questions about G-Forces on pilots.

The article says:

  The average, fit aviator has about a 3g "naked" limit, which comes
  from automatic tightening of blood vessels forcing blood to the brain.

I'm a private pilot, with no training with G-forces.  I'm average fitness (maybe even a little below average).

For the big 40th Birthday, I went to the Sky Combat Ace experience, doing simulated dog-fights in Extra 330LCs.

The instructors said that both my brother and I pulled in excess of 7G. (with no G-suits or anything beyond some grunting).

I'm not so sure I believe that now.  Were they blowing smoke to give us a fun experience?
Or could I have credibly pulled 7G for 10-20 seconds without prior experience?

The  G-limit for the Extra 300 (the parent airplane to the 330):

  The Extra 300 is stressed for ±10 G with one person on board and ±8 G with two

So it is certainly possible from an airplane perspective that you pulled that many Gs - in an aerobatic airplane, it's not unheard of to pull 1G less than the aircraft limit (usually there is a large margin of safety built into that number before a catastrophic structural failure).

Now did you pull 7Gs? As others have noted, 7Gs for 10-20 seconds seems very unlikely. But it is not out of the question your peak instantaneous G load was 7.

From a human factor standpoint, G-suits typically don't increase the max G-load, but rather increase the sustained G-load and the time period it can be held for. G-suits do very little to actually decrease blood flow (pooling in the legs) - they primarily serve as a reminder to the pilot to perform an AGSM.

  A g-suit does not so much increase the g-threshold, but makes it possible to sustain high g longer without excessive physical fatigue. The resting g-tolerance of a typical person is anywhere from 3-5 g depending on the person. A g-suit will typically add 1 g of tolerance to that limit.

Now there are a vast array of factors affecting this resting G-tolerance. I flew  a military high-performance trainer, and my resting G-tolerance was around 3, or the lower end of the spectrum. One of the other pilots I knew had a resting G-tolerance of about 7.5. The primary difference is blood pressure. I am 5'10" and only 145 pounds, with a resting blood pressure of about 90/60-100/70. He was 6'0" and about 215 pounds, with a much higher blood pressure. Tall, skinny, nonsmoking runners have actually the worst resting G-tolerance while short, stocky, sedentary or heavily muscled smokers have much higher tolerance. Other factors, such as how hydrated and rested you are and how much stress you are feeling, can vary vastly your daily G-tolerance.

All of those numbers I just listed do not take into account the pilot performing an AGSM (Anti-G Straining Maneuver). This can add several Gs to the pilot's tolerance, and is one of the biggest things taught in physiological training for high-G maneuvers.

So TL; DR: could you have pulled 7Gs without prior exposure? It is possible, depending on varying personal factors and body type, but it is not likely. However, with a little training, 7Gs is perfectly attainable.


What makes the Air Force and Navy pilots different? As far as I know, aviation in both divisions are almost the same. They fly various aircrafts, they do dogfights, air to ground attacks, air transports, electronic attacks, and more. Can you say that Air Force and Navy pilots are different in some ways?

Navel aviation pilots are specifically trained for anti ship and anti submarine warfare roles and they are also trained for carrier based operations. They are also trained for ground attack, dogfight etc. Air Force pilots does not have carrier based landing skills and they do not have anti submarine warfare capabilities.


Why not leave an electric fuel pump on all the time? This question is related to single-engined piston airplanes equipped with an electric fuel pump like the Piper Dakota/Arrow etc. The POH states that the fuel pump should be switched on for takeoffs/landing/approach and maneuvers.

I understand that the fuel pump is a backup system and switching it on provides an extra margin of safety in such situations. My question is, why does the POH recommend that we switch off the fuel pump after reaching a safe altitude (say 1000 ft). Wouldn't it be safer to just leave it on? Is there potential for harm in leaving the electric fuel pump on? Or this is done on common sense grounds: just to save maintenance and prolong the life of the backup system so that it is there when you really need it?

The electric fuel pump is used to provide an extra safeguard against fuel starvation during critical operations in siphon fed fuel systems where either there is an increased risk of this happening or where sufficient altitude does not exist to resolve a problem such has switching fuel tanks or during takeoff and landing operations.  While one could in theory continue to operate the electric fuel pump continuously throughout flight operations, it it not necessary and just adds additional wear and tear on the unit.


E class airspace(transition area) Extension area(E airspace) shown as shaded magenta start from 700AGL but how to measure 700AGL over the water like in this picture?
Cuz there’s no land…

While the documentation uses AGL when talking about class E airspace the chart legend uses the slightly different phraseology of

  above surface level

This avoids the ambiguity of, ground or ocean.

Interestingly the glossary uses both ground level and sea level but explicitly defines neither.


Why do most airliners have underwing engines, while business jets have rear-mounted engines? Why have modern airliners converged on a configuration with two or occasionally four underwing engines, while modern business jets have converged on a configuration with two or occasionally three rear-mounted engines?

I am of course aware of the many past airliners with rear-mounted engines, with the DC-9 and MD-80 and later derivatives continuing into production until rather recently, but it still seems fair to say that the underwing configuration dominates modern airliner design.

Meanwhile business jets exclusively use the rear-engine design, unless one counts small airliners put into business service such as the Boeing Business Jet.  Does the underwing engine design simply not scale down well to something the size of a smaller business jet?

Here's a related question with good answers but still doesn't really answer my question-- How does the mounting location of a jet engine affect aircraft performance?

One key factor now occurring to me is that business jets are boarded with self-contained airstair doors/ steps so a fuselage sitting high off the ground would be problematic.  Airliners with long fuselages perhaps may need rather long landing gear to provide adequate ground clearance for rotation anyway-- although the DC-9 / MD-80 with its wing and landing gear far to the rear did sit rather low to the ground-- still in the airliner situation perhaps there simply is not as much benefit to having a low-sitting fuselage as there is in the business jet situation.

Of peripheral interest is this Wikipedia article on "airstairs", pointing out some cases where full-blown airliners are equipped with airstairs.  One of the more interesting lines is:

  The most unusual airstair design was found on the Lockheed L-1011,
  which was a full-height airstair which was stored in a cargo
  compartment and allowed access from the right aft passenger door to
  the ground. This design was ultimately so large and heavy, and it took
  up valuable cargo space, that it was rarely used.

Underwing engines have more advantages for larger planes, but there must be enough room underneath the wing to mount them.

This was tested by Boeing in the 1960s, when they designed the B737 and had two competing teams pre-design both configurations. The underwing config won. When Airbus designed the A320 two decades later, the underwing config won again with engines that were a good amount wider.

Image source.

Advantages of underwing mount:

Engines mounted in pods underneath the wing reduce the upward bending moment of lift creation, and therefore allows for lighter construction.
If the pods stick out forward of the wing, the backwards aerodynamic twisting moment is also reduced: lighter construction.
The engines can be easily reached for maintenance.
All of the fuselage can be used for useful load - tail mounted engines have structural reinforcements in the fuselage between the engines.

Advantages of fuselage mount:

The yawing moment of a single engine fail is smaller, allowing for a smaller vertical tail.
Fuselage can be lower to the ground and allow for low door sill height.
Landing gear can therefore be shorter.
Clean wing configuration.
Lower trim change due to thrust.
Less noise in most of the cabin.

In smaller planes, the wing bending & twisting moment is a relatively smaller constructive issue than in larger planes. And indeed, being closer to the ground allows for easier boarding via air stairs.


How does an aircraft distribute fuel equally between powering the turbines on the engine and injecting it into the afterburner? How does an aircraft distribute its fuel equally between powering the engine turbines and injecting it into the afterburner?

The aircraft fuel system does not  distribute fuel equally among the afterburner and the engine's core (i.e. the compressor-combustion chamber-turbine part). It distributes fuel in such a way that the combustion is maintained in both the combustion chamber and afterburner and the required thrust is produced.

Consider the case of F-15 E/I/S or F-16 C/D powered by PW F100-PW-229 (data from jet-engine.net). Here, the thrust and SFC (specific fuel consumption) for dry and wet thrust are given as 17800/29100 lb and 0.726/2.060 lb/lb/hr respectively.

Simple calculation yields fuel consumption rate of 59946 lb/hr and 12923 lb/hr with and without afterburners. This answer by @KeithS gives the fuel consumption of F-16 with and without afterburner as 64000 and 8000 lb/hr respectively. Also note that fuel consumption varies greatly with other conditions like altitude, external stores etc.

It can easily be seen that the use of afterburner increases fuel consumption by 5-8 times and most of this has to go into the afterburner. So, once the pilot punches the afterburner, the fuel system delivers much more fuel to the afterburners than to the 'core' of the engine. These decisions are made usually by the aircraft computer/engine FADEC based on pilot input in all modern combat aircraft.


If the landing gear is stuck in an airliner, can the crew go down and manually open it? There are several instances where we see the landing gear getting stuck before landing. In these kind of situations, can the crew, or one of the pilots, go down to the landing gear area and solve the issue?

On older (WWII-era) large planes, the crew had access to a manual crank to lower the gear if the automatic extension/retraction system (usually a linear actuator powered electrically) failed. I don't know of any that had a gravity-based manual system; I'm sure a few bomber crews would have appreciated something like that after having to crank the handle to get a disabled main gear locked. Here are the manual crank ports for the B-17, on the forward bomb bay bulkhead with red arrows indicating the direction to crank:

The B-17 had a retractable tail gear as well, so there was a third manual crank in the tunnel to the rear gunner's position (but if there were multiple gear failures, the main gear would be the priority).

On more modern airliners with hydraulic extension/retraction, manually cranking the gear into place isn't as much of an option (especially since any access to a crank linkage would be in either the passenger cabin or the cargo hold). For these aircraft, the gear are designed with a "manual extension" or "gravity release". This basically disconnects the gear strut from its uplocks and the hydraulic cylinder, allowing gravity to pull the gear down and then momentum (or a spring mechanism) will engage the downlocks. Here is a diagram of the 737's manual gear panel on the right-hand wall of the flight deck:

The 747 has a manual nose gear extension system in the avionics compartment below the first-class cabin. The 747's main gear are much more complex, but it is still possible to let gravity pull them into place, by using a set of "alternate gear extension" switches on the overhead panel to release each gear from the hydraulics and uplocks:

It doesn't always work; a Virgin Atlantic 747 pilot was forced to make an emergency landing with the outboard right main gear stuck in a half-extended position:


What safety requirements are there when designing an electronic device to go inside the cabin of an airliner? I am considering designing and building an automatic air-sampling unit that can be placed in the cabin of an airliner and left unattended for an extended period of time.  The device will use an electronic air sensor and (small) processor so that it can detect and sample fume events without crew intervention, will be battery powered using a small battery of a known-reliable type (i.e. not Lithium-Ion :), will not have any functions that intentionally emit RF, and will be packaged in an enclosure (either plastic with internal shielding or metal).

What EMC and electrical-safety requirements/specifications must this device meet in order to avoid posing a hazard to the aircraft it is placed on, and how do I ensure that this box is not mistakenly considered a security threat to the aircraft?

A document called RTCA DO-160 defines the testing for a number of environmental requirements. Within each test standard is normally a number of different test conditions. The level of required testing is normally defined by the customer of the device, who has the task of showing the regulating body (FAA, EASA, etc) that this is sufficient to ensure the aircraft meets all regulatory requirements.
In addition, DO-254 gives guidance for electronic hardware and DO-178 handles software. There are other legacy documents like MIL-specs that define testing as well.


Are there technical means to enforce aircraft speed limits as well as regulatory means? Do any aircraft have a device which can be pre-set or can be controlled from the ground in order to prevent the aircraft from exceeding a speed limit?

These could be similar to what they have for buses in many places so the bus driver can't go beyond a speed limit.

Some aircraft have flight envelope protections which, among other things, aim to prevent the aircraft exceeding a predefined maximum airspeed.

High Speed Protection:
Prevents exceeding VMO or MMO by introducing a pitch up load factor demand.
The pilot can NOT override the pitch up command.

- AirbusDriver.net
Note that these are structural safety limits, people on the ground cannot use this feature to impose arbitrary speed limits by geographic area, time of day etc.


Is there any theoretical problem powering the fan with an electric motor A question that I have been wondering about lately is whether it would be possible to power the fan in a modern hi-bypass turbofan with an electric motor and achieve performance comparable to the kerosene-fueled engine. Or is the 10% of air that flows through the combustion chamber crucial for the operation of the system?
(I wish to ignore all things that make electric aircraft unviable, like power-to-weight ratio of engines and battery capacity)
Edit: I have realized that the "electric" part of the question is really not the core of the question. What I am really interested in is whether it would be feasible to power a plane using only the fan and not the exhaust jet with a performance similar to a current jet (which would make the engine similar to a turboprop with many blades and a cover around it)

Taking into account your wish to ignore all things that make electric aircraft unviable, like power-to-weight ratio of engines and battery capacity, the answer is yes.
You would, however, need to compensate the missing jet engine thrust by increasing the fan thrust.
Otherwise the answer is no.


Is it legal to install a TACAN receiver on a civilian aircraft? It is legal to use the DME from a VORTAC station (see Is it possible for a civilian to use a VORTAC station? for details), but is it also legal to use the TACAN for bearing determination?

This would probably require different equipment, which is not typically installed on a civilian aircraft (see Why doesn't civil navigation equipment support TACAN? for details). But the answers to that question do not mention whether or not it is legal to install and use this equipment.

You can buy and legally use TACAN equipment for GA aircraft from Bendix/King, but the system is expensive (close to US\$100,000), relatively heavy and requires rack-mounting.


Advantages to doing an initial instrument rating training in a Cirrus over a conventional aircraft? I live in Australia and I'm in the process of choosing a school for my instrument rating. I need feedback from the many of you out there with IFR experience willing to share their thoughts.

I'm going to do a single engine instrument rating because I am a just private pilot with no immediate plans of doing a CPL or flying multi engine aircraft at the moment. I have CSU and RU endorsements, have flown VFR to this day on a basic and complex aircraft. There are a few schools that I'm considering, with the only real difference among them boiling down to aircraft and, of course, total training costs.

1. Towards the top end, I could train on an SR20 and brand new full
    motion simulator in a posh school.

2. The middle range would be on an C182 or a PA28R and standard
    simulator.

3. The lower range would be on an IFR rated C152 + time on a complex
    aircraft, and again standard simulator.

All up, the costs would go from simple on the C152 to (almost) double on the SR20.

Here's my call for advice: Other than the fanciness of flying in an SR20 over a C152 and learning in a shiny school,

What are the real, concrete benefits of doing an IR in an SR20 over a more conventional aircraft?

On the one hand, the C152 has analog instruments while the SR20 is all glass and has autopilot which, I understand, comes down to personal preference. Another key difference is speed. I'd say that flying more slowly might be an advantage to stay "ahead of the aircraft" but then again, that might be beside the point of IR training. I'm the super lucky position to be able to afford both but I still want to make a financially clever decision as I'm not the kind to spend money unnecessarily. In hindsight, that's how I did my PPL and it served me well.

A great many thanks for your much valued answers.

PS. Another factor that does play a role is obviously instructor experience and their availability.

PPS. A definite minus for the SR20 is the "requirement" to do the CSIP endorsement which puts the total costs just shy of that of a multi-engine endorsement on the Seminole or on the DA42 + IR.

There aren't any. Unless you are planning on purchasing a Cirrus, you'd just be burning money for no good reason by training in that. An instrument rating is an instrument rating regardless on what aircraft you train. What is important is the quality of the training, ground and flight. If you think about it, I am quite sure you can find better ways of spending the money you will save.


What are the different variants of ILS intrumentation? Background: I'm a huge aviation enthusiast with a lapsed PPL. Have flown some larger aircraft on MSFS, but never flown anything in real life bigger than a C172.

I was lucky enough to do a bit of Simulator time on a B737 simulator this weekend. One thing I wanted to try out was an ILS approach in very limited visibility. The instructor set me up, and asked what ILS display I was "used to seeing". I described it, and she set it to this - totally what I was expecting:

(Thats the purple vertical bar for centre line and purple blip for glide slope).

So, my question is, what was the "Other" option? This may be specific for a B737, or it may not. Why are there 2 types of instrumentation for an ILS approach? Is one a newer system, and is it therefore preferred.

In the Boeing next-generation "NG" flight decks, the PFD can be configured one of two ways. The first way, is the default presentation, where you only see to top of the HSI. The pilot will fly the aircraft via the pips along the bottom and right of the attitude indication (and usually with the flight director cross as well):

If my memory is correct, Boeing, NASA, and others did a number of human factors studies over the years and found that this layout allowed pilots to more quickly understand critical flight information, and this layout is now standard even in GA glass cockpits (e.g. Garmin G1000, Collins Proline).

The second configuration provides a virtual "six-pack":

This configuration is primarily used for airlines that are transitioning between "Classic" and "NG" flight decks. One airline I remember in particular is Southwest, as they had a combination of 737-300/400 and -700/800 aircraft and having a classic configuration on their new flight decks allowed all of their pilots to fly any of the aircraft, without needing significant differences training. The particular presentation can be customized per the airline's needs. An airline will typically move exclusively to the modern PFD display once all of the older aircraft are moved out of the fleet.


Can you identify this plane from 1920 and prior vintage? My grand uncle, that is my grandfather's brother had a flight school and performed in barnstorming shows somewhere in the midwest and it's not a Jenny. His name was Ralph Bloxham and is seen with his flying cap on in the rear cockpit.

I've found a few planes with similar characteristics but none with them all. Can you identify the plane?

Single bay biplane.
N- type interplane struts.
Little to no dihedral.
Rear cockpit is just below the trailing edge of the top wing, no cutout in the top wing.
Dual (upper and lower wing) ailerons with push/pull link.
Non-radial (inline or V engine, likely water-cooled but possibly air-cooled such as a Franklin)
Two blade prop.
Bipod cabane struts at the front spar, single rear spar cabane strut.
Mild stagger of the top wing forward, Wire wheels (may be missing covers)
No brakes, interesting side panel on fuselage with a radiused front edge, rather long- coupled empennage which is not visible in the photo.
Fuselage is rather slab sided with pretty constant taper and is very slender aft of the cockpits.
Bottom of engine cowl tapers up toward the front, and no protruding push rods or exhaust manifolds visible.
Engine cover is not higher than the fuselage contour line (no hump)
Exhaust out the side or both sides, don't know but there's some semblance to the liberty engine.

The N strut design seems to be a rather uncommon trait as I search, most are wired only, and so is the completely enclosed long engine cowl/cover which rules out all radial or rotary types. Single bay rules out many. What do you think?

I'm pretty sure that this is a Lincoln PT, part of the Lincoln Page PT series, although the PT/PT-K was in the late 20's and early 30's:

Source: Wikimedia Commons

You can read about it from EAA:

  The Lincoln PT-K model was the next development of the Lincoln-Page PT, designed specifically as a flight training or low cost sport flying airplane. The “Page” had been dropped from the company name before the PT-K came out, so the “PT” stood for “Page Trainer”.

  The PT-K was considered a well designed airplane from an aerodynamic standpoint in 1929, taking into account the mannerisms of a student pilot. The fuselage had a long moment arm for the tail surfaces, meant to desensitize the pitching and bucking of a nervous student pilot as he worked his way through the curriculum. In other words, the Lincoln was not too fussy and wouldn’t get all upset due to a little pilot error.

  The PT-K was powered by a 100 hp Kinner K5, which could max out at 104 mph. It was no exceptional trick for the PT-K to perform a well rounded loop from level flight; there was really no need to dive and gain excessive speed to bring it around.

  The structure of the PT series proved to be quite rugged. The original design had a split axle landing gear configuration with a shock absorbing bungee chord. The later design featured a stiff-legged rigid gear, coupled with low pressure airwheels. The low pressure airwheels had more than enough “give” in the sidewall so that no further shock absorption was needed.

  The PT-K was available with a few options, including a metal propeller, a lightly framed fully enclosed canopy for cold weather flying, navigation lights, and engine starter. One other item available as an option was Bloxham “Safety Sticks”. This option allowed the instructor to disengage the student’s control stick if a student “froze” during the flight.

The aircraft in the picture is restored, but the strut matches especially around the cockpit.

The Lincoln PT was offered with a different V-8 engine, the Curtiss OX-5 but there is no evidence that one was ever sold. You sir may have that evidence as you don't seem to have the PT-K version, but the PT version.