Which altitudes are safest for VFR? As a fairly new pilot who does a fair amount of x-country flying I find myself "feeling" safer going East at 5500 and West at 6500 happily burning extra fuel to get up there. Listening to the on route frequency position reports as I fly, I find the vast majority of GA traffic fly between 3000 and 4500 giving me the feeling that things are pretty crowded and less safe down there. Is there any validity to this?

The altitude you fly at should be determined by the following:

Where you are safest.  Crossing 5500 foot mountains at 6500 simply is not a good safe practice.  Your altitude should be higher at night, if for nothing else, to give you more time to diagnose a problem or pick a suitable landing spot.  When my primary students fly the hills around here, I insist that their final cruise be above 3000 AGL.  That keeps them out of 2000 foot towers (most are just 1500 feet), and gives them time to pick a landing spot, and gives them altitude for getting radio assistance from FSS or ATC.  Higher is better, but I tend to not encourage 10,000 foot level flights for student pilots because there are other risks.  You should consider all the risks associated with altitude.  Lumbering along at 11,000 feet may not be the least risk option.
Your altitude selection should help you in navigating.  Being up higher lets you pick landmarks, see more VORs, reach more ATC and FSS facilities.  Generally, my students do not use GPS (as student pilots), so if they get lost, they have to descend and read water towers and things like that to find themselves.  That doesn't happen often.
Your altitude should keep you in good weather.  If you are a VFR pilot, you will probably not want to make your trip on top, with a 1000 foot ceiling below you. Again one needs to balance risk.  Being at an altitude which keeps you below the scattered layer, keeps visibility good, and can make a trip more enjoyable.  But if Nervous Nancy is taking her first plane trip, and is predisposed to motion sickness or is simply nervous about bumps, then you want to pick an altitude that will be a bit smoother.

The whole exercise of picking the right altitude, the right route and even the right day is part of a big optimization problem, where you are looking at safety, enjoyment, economics and other factors.  There is no magic altitude, no perfect route.  What is good 80% of the time may be not so good 20% of the time, which is why ADM (Aeronautical Decision Making) is emphasized in training and recurrent training.  Having good visibility, nice tailwinds, a smooth ride and plenty of altitude over the terrain are always nice.

Oh, one exercise you might try...I tend to do this with flight review students...we do airwork at altitude over the hills or low mountains about 25 to 30 miles from the airport.  I pull power and they dive for the nearest pasture, which has down slopes which exceed the descent profile of the aircraft with power off, the gear and flaps down, and heavy slipping.  Or select a field which has an up slope which is extreme enough so that a landing is impossible.  Then I suggest that they try the home airport.  "But it's 20 miles away."  or maybe 25, and they might have a good tailwind at 5000 feet above the airport.  And they start the glide.  On the way we pick alternates, including flat grass strips, along the way.  Almost 100% of the time, the lightly loaded aircraft makes the airport, and they are amazed that they were able to travel so far without power.  And the take-aways are many, which include that having a couple of thousand feet below you really opens up your options.  And to not pick the first field you see.  With altitude, there may be others as you get to more favorable terrain.

I have managed to fly many decades, many hours and never so much as scratched paint on an airplane.  However, before every flight and during every decision, I ask myself what an objective, after the fact, review board would say.


How should an aviation novice evaluate 'flight experience' or 'flight training' services? Some companies offer a service wherein a complete novice may take a flight in a small aircraft, sometimes as a passenger, sometimes 'taking the controls' under the close supervision of an instructor. Examples of this include:

Spitfire Flight Experiences
Fly a MiG-29
Laser dogfighting

How should a novice, who has little or likely no flight experience at all, evaluate whether an offer such as this is:

Safe, within reasonable bounds (ideally the same safety precautions that a well-run national aviation regulator would require for full-time student pilots)
Legal, again within the remit of the regulations that would apply in that particular country
Reasonably likely to offer what is claimed i.e. where 'taking the controls' is advertised, that this would constitute actually manoeuvring the aircraft and that this would be legal and proper within the regulations

For Laser dogfighting, you linked to SkyCombatAce.
I did the Sky Combat Ace experience in Las Vegas about 4 years ago, and there were no lasers involved.  Rather, the back-seat instructor judged when you were in position for a kill shot, and just radio'd "kill", at which point the other plane activated smoke and started spinning.
Although I enjoyed my experience with them, its worth noting that they'd have 2 major crashes resulting in 4 deaths in about the last two years.   The investigations are still pending, but personally, I don't think it looks good.
In the April 2016 crash, the local weather was announcing thunderstorms in the area, and the 3 planes took off anyway, initially requesting flight directly towards the storms until the Tower Controller suggested they reconsider.  They either seemed oblivious to the local weather, or didn't care about it.  It has not yet been determined if that was a factor in the crash.
On the one hand, I enjoyed my flight.  On the other hand, I now have serious reservations about their operation, and would not consider flying with them again.

As far as the "taking the controls" aspect:  I had my Private Pilot's license at the time I flew with them, so I'm familiar with flying, although not at the performance levels of those planes.  However, "taking the controls" meant the ailerons and elevator only; no throttle and no rudder (controlled by the instructor).  In that sense, it was a lot like a simplistic video game with a joystick.

Update: July, 2018
The Factual Report on the first accident, in Henderson, NV is now available on the NTSB website.  The report took over two years from the incident to come out, which is atypical (6 months to 1 year is more common).
The report had two concerning findings:

First and foremost, the report finds the pilot had detectable levels of THC (marijuana) in his blood and tissue.  The report points out that it is impossible to tell how recently the pilot used it or if he was impaired at all. However, that he had any amount of an illegal substance (marijuana did not become legal in Nevada until several months after the accident) is deeply disturbing.

Secondly, the report calls into question if the flight operation is operated legally by CFR 14.91 and .61 (Instructional Flights), and suggests they may have been legally required to hold a part 119 air carrier / commercial operator certificate and operate under the stricter 121 or 135 rules.

I will reiterate that I enjoyed myself during my flight with SCA.  But after two fatal accidents and this incredibly concerning report, I cannot recommend that anyone fly with them until major legal and operational issues are resolved.

Update: May, 2021
A GoPro camera which captured much of the second crash was found after two years.
The additional information it helped reveal is disturbing.  A detailed article is here.


Is a chambered fan more efficient than a ducted fan? A chambered fan is not exposed to the oncoming airflow in forward flight.
I am working on a VTOL design that uses a chambered fan.

There will be a wing attached to top of the thrust arms.

Could it produce 12,000 pounds of lift force using a GE T901 with 3,000shp with 8' dia. contra rotating fan?

The design is not very efficient:

It does not use forward airspeed to pre-compress the airflow entering the engine.
It blows exhaust air into the air inlet.
It bends the exhaust flow 180 deg, creating friction and loss of thrust.

Short, straight, unimpeded airflow works best.


What is the chord of a wing? In the answer to this question the author mentions the chord of a wing.  What exactly is that, and how is it used in wing design?

the Chord is the line stretching from the leading edge of the wing to the trailing edge parallel to the centerline.

There are different variations to describe the chord of the whole wing.

The Standard Mean Chord
Defined as $\frac S b$ with $S$ the surface area and $b$ the span of the wing. This is not used in aerodynamics.
The Mean Aerodynamic Chord
Defined as $\frac2 S \int_0^{\frac b 2} c(y)^2 dy$ with $y$ the coordinate along the wing and $c(y)$ the chord at $y$.
For swept or delta wings this gives the equivalent chord for a rectangular wing, and is important for placing the Center of Gravity for stability.


What is a 5 Tube System? I've been going over the new Master Minimum Equipment List (MMEL) for Textron Aviation Model 500 Series Citations and on page 70 an Electronic Flight Information System (EFIS) 5 or 3 Tube System is mentioned but I can't find any information or pictures of it online.

Can someone explain?

Look for the Honeywell Primus 1000 System.

Early systems were 5 tube - an EADI and EHSI for each crew plus an MFD (Nav/Wx display) as shown below:

source: swaviation.com

Later ones were 3 tube - 1 PFD on each side plus the MFD as shown below:

source: airliners.net

The symbol generators are shown in the MMEL as they have reversion capability. The 'tubes' or the actual displays would be listed under section 31 Indicating/Recording Systems but aren't included in the MMEL as they are always required.

Section 34 Navigation does allow for functional portions of the displays to be inop.


G1000: Is it possible to balance the radio volume for pilots and passengers? I regularly fly a DA42 4-seater equipped with a G1000 system. Sometimes my passengers in the rear seats complain that the radio volume is too loud for them but it is OK for me (intercom is just right for everybody). Reason probably is that I am using a very different headset than them which I assume has a higher impedance and thus requires more electrical power for the same audio volume. If I turned down the COM volume it would be too quiet for me.

I wonder if there is any trick to adjusting/balancing the volume in such a setup.

You can't control the radio volume separately for the passengers but you can control the intercom volume separately. If the passengers' headsets have individual volume controls you can ask them to turn down the volume and you can then compensate by turning up the passenger volume for the intercom. That way the overall intercom volume will be about the same but the radio/ATC volume will be lowered.

If all else fails, you could turn off the radio completely for your passengers :-)


Will I fail a skills test for doing a constant-aspect climb-out? The climb-out for one of the runways at my local aerodrome is over a residential area, with fields to the left. My instructor has had a bee in his bonnet lately about engine failure after take-off if the take-off run is longer than it should be (e.g. after a touch-and-go). He's started encouraging all of us to perform constant-aspect climb-out into the circuit, where we start the turn onto the crosswind leg at the aerodrome boundary regardless of height, and keep the runway end at the same declination in our field of view as the turn progresses. This means we point towards the fields much earlier on, but it hurts the climb performance to do it.

Although I can see the wisdom of being ready for EFATO, I'm concerned about practising a non-standard circuit which contravenes the Noise Abatement Procedure. (The NAP forbids turns below 500 ft, although it is overridden by measures to ensure flight safety.) When I come to do my PPL skills test, would this unusual circuit be acceptable, or will I have to fly a standard circuit with a straight climb-out, turning onto base at 500 ft? I'm in the UK and studying towards an EASA PPL(A).

I fly in the UK, and while there is certainly a "standard circuit" (ie pattern) in the syllabus there are many, many fields that have a non-standard circuit for one reason or another. This is often to avoid antagonizing unfriendly neighbors.

There's also the fact that the NAP is not a rule or regulation, it's more how the airfield management would like you to fly.

I personally don't like the constant aspect climb-out idea at all, it will mean you are paying more attention to aspect than airspeed and looking out for other traffic. It also means that the pattern traffic is not predictable for other visiting your airfield or by others in the pattern. So I would push back on your instructor on this.

As for whether you will fail your skills test if you do it would depend on the examiner's view, and what they are looking for. While a constant aspect departure shows skill, I personally would take a dim view of people doing it despite the published departure procedure, even if it is non-binding. I would discuss it with your examiner ahead of time, and practice your circuits how they say they want it, or at least discuss it with them on the day.


Why does the flight controls check come before arming the autobrake on the A320? As per SOP on the A320 we first perform the flight controls check and then arm the autobrake to max. What is the reason behind this?

Both are in taxiing SOP.

The reason is in case the spoilers were accidentally left extended on the flight before (or during a maintenance check before the flight by the line engineers).

If that's the case, then pressing MAX while taxiing before checking the flight controls (spoilers included) will engage the full RTO brakes and surprise the crew, injure a cabin crew member, and/or cause an accident (a ground vehicle thinking you won't stop and is aiming to pass behind but then you stop in an instant).

Source: Operational Liaison Meeting Taxi Procedures (PDF page 7).

I'm guessing someone has done it before.


At what Mach Number do variable cycle engines (VCEs) start showing benefits over turbofans? Currently, most supersonic engines engines are low bypass turbofans. However turbofans start becoming infeasible as the mach number increases. Variable Cycle Engines, still have noise and weight issues but are purported to give significant improvements in spillage, airflow matching and mixed mission cycle efficiency.

I wanted to know what is the lowest mach number (roughly) at which the benefits of the VCE's attributes are purported to give it an advantage over turbofans.

I'm mostly interested in benefits in specific fuel consumption thrust, spillage and matching losses etc. and am not that worried about the noise, weight and complexity problems (though useful information regarding this is appreciated).

most literature regarding this that I found were from the 1960's to the 80's so links to newer resources, literature, etc. are appreciated

It's not a high Mach number that gives a variable cycle engine an advantage over a turbofan. Most air-to-air combat happens at subsonic speeds, because drag gets very high when you start trying to turn sharply at supersonic speeds. For combat, you need high specific excess power (SEP), which comes from a low bypass ratio turbojet. But, cruise speeds are also typically 0.8 to 0.9 Mach, even for fighter jets. But lower fuel consumption in cruise occurs with a high bypass ratio turbofan. That's because the exit velocity is not much higher than the aircraft velocity, which means the propulsive efficiency is high. But, it means the aircraft doesn't have excess power, to suddenly accelerate, or stop energy bleeding off in multiple, rapid turns.

So, in both cases, cruise or combat, speeds are subsonic, but the desirable type of engine is quite different. To illustrate, the specific fuel consumption (SFC) of a CFM56-7 is about 0.37 lb/hr/lb, with a BPR of about 5.4. In comparison, the F100-PW-229 has an SFC of about 0.73 lb/hr/lb (non AB), and a BPR of about 0.36. That's pretty close to a factor of two. Obviously, a VCE engine can't change the bypass ratio as much as in this example. I don't expect the variation of bypass ratio (BPR) in a variable cycle engine is published anywhere.

So, the advantage of a VCE occurs even if you don't need or want to fly supersonic. The advantage occurs when you suddenly need excess power, even at subsonic speeds. Military aircraft benefit from a variable cycle (variable BPR) not because of the large speed range over which they fly, but because of the wide variety of mission types they fly. Civil aircraft just need to be optimised for cruise. Variable cycle allows military aircraft to be better optimised for combat and cruise, not just the first.


What are these "F,S,O" characteristic speeds? I am leaning for an avionics exam and I encountered a question asking for the meaning of F, S and O . They appear to be "Characteristic speeds", which are calculated by the flight management system (FMS):

F speed is the minimum speed at which flaps should be retracted from CONF 3 or 2 to CONF 1+F.
S speed is the minimum slats retraction speed, i.e. the minimum speed at which a clean configuration should be selected.
GD speed is the engine-out operating speed in clean configuration. In other words, it corresponds to the speed that allows the highest climb gradient with one engine inoperative in clean configuration.

source:
https://www.icao.int/WACAF/Documents/Meetings/2015/ICAO-BEA/SafetyFirstissue20.pdf


Is it possible to celebrate a holiday twice by flying from one point on Earth to another? It recently came to my attention that the different time zones on Earth could allow for a person to celebrate holidays such as New Years twice by flying from one point to another.

Is it possible to fly quickly enough to celebrate a holiday twice?

Technically, yes. As I right this, it is half past midnight at London, Heathrow. It is half past eight at New York Kennedy. And, it is half past five at Los Angeles. If it were not Daylight Savings Time, it would be half past seven in New York. On New Years Eve, you would have between 4 and 5 hours to celebrate in London, fly and land, before you started celebrating again In New York. Hardly any time for even a celebratory drink (as a passenger) and kiss. I have come very close to landing before I took off  when traveling from the US to the Asia. So far, I have never made it. I’ve always missed it by an hour or three.

It would be easier to celebrate New Years in somewhere like Lubbock, TX (Central Time, USA), then flying to Roswell or Clovis, NM (Mountain Time, USA).


What factors are included in the moment coefficient of a wing? I am trying to make sure I correctly understand the moment coefficient. Both the lift and the drag coefficients vary with angle of attack and can be either positive, negative or zero. The moment coefficient pertains to the moment specifically due to the aerodynamics force (lift force on the wing mostly). The moment of a force can be calculated about any arbitrary point on the chord (or even outside of it). For a positive cambered wing, if the pole was the arbitrary point X, as the angle of attack varies (positive), the moment would vary and be negative.Is that correct? If the angle of attack goes negative, can the moment coefficient go positive?

If the pole X is chosen to be the aerodynamic center AC (25% of the chord distance from leading edge), the moment of the lift force calculated about the AC remains instead constant with varying the angle of attack.

For a positive cambered wing, I understand that there is a moment due to the lift force (mentioned above) and also another intrinsic nose-down moment (which I think it is constant for all angles of attack) solely due to the cambered shape of the wing. This intrinsic moment is nonzero when the lift is nonzero or even zero. Does the value of the moment coefficient take this intrinsic moment into account as well besides the moment due to the lift force?

Thanks.
Brett

As much as I try to explain aerodynamics without resorting to potential theory, sometimes it is indeed helpful. If you interpret the pressure distribution around a wing as the sum of

a zero angle of attack distribution due to camber and
the Birnbaum pressure distribution due to angle of attack,

things should become clearer.
On the left I show a cambered airfoil at zero angle of attack and on the right an uncambered airfoil at different angles of attack. The first cambered airfoil is a Joukowsky-type which has an arc section for its camber line. The one below it has an irregular camber line that is simply the center between its two surfaces. If you draw circles inside the airfoil at different chord locations, the camber line is the connection of the centers of those circles. The higher the local curvature of the camber line, the higher the local lift at that chord station.

The remarkable thing about the uncambered airfoil is that its lift distribution over chord is geometrically similar between different angles of attack. It only scales in magnitude, but keeps its general shape. Therefore, its center of pressure is at the quarter point for all angles.
The lift due to camber does not change with angle of attack and can be added to the lift of the symmetric airfoil in order to arrive at the total lift for a cambered airfoil at a nonzero angle of attack. Since the lift due to camber dominates the total lift at small angles of attack and becomes more and more insignificant as the angle of attack increases, the center of pressure will move from the point due to camber to the quarter point as the angle of attack rises.
If the lift is composed of a force acting at the quarter point plus an offset moment, this moment is only determined by the camber lift. Since that is constant over angle of attack, so is the offset moment.
This is all strictly true only in inviscid 2D flow, but gives a good approximation for real wings with enough aspect ratio and no flow separation. Viscosity will muddle the picture a little and with smaller aspect ratio the center of pressure of the symmetric airfoil will move forward, but for most airplanes the above captures reality rather well.

For a positive cambered wing, I understand that there is a moment due to the lift force (mentioned above) and also another intrinsic nose-down moment (which I think it is constant for all angles of attack) solely due to the cambered shape of the wing.

Only if you decide to let your lift force attack outside of the quarter point. Then the lift moment caused by the distance between the force vector and the quarter point will produce that moment due to the lift force. If, however, the lift attacks in the quarter point, only the intrinsic nose-down moment which is constant for all angles of attack will be left.

This intrinsic moment is nonzero when the lift is nonzero or even zero.

Yes, indeed. At the right negative angle of attack, the downforce from that angle of attack will exactly compensate the camber lift, leaving the whole airfoil at zero lift but still with the constant moment.

Does the value of the moment coefficient take this intrinsic moment into account as well besides the moment due to the lift force?

Again, only if you (against all convention) decide to define your lift force in a point different from the quarter point. As the moment coefficient is defined, it contains only the "intrinsic moment".
When separation starts (and even with a thicker boundary layer) this will gradually break down and the moment will fluctuate. But then we move outside of the linear range of aerodynamics and potential theory will no longer apply.


Why does the stagnation point on airfoil move with the angle of attack? What is the physical explanation for the reason the stagnation point is where it is? And, why does it move further down the lower side of the airfoil with an increase in angle of attack?

At the stagnation point, the streamline is directly perpendicular to the airfoil. As angle of attack increases,  the streamline is no longer perpendicular to the leading edge (which is where the stagnation point is usually located at 0 angle of attack). This is the best explanation I can come up with for why the front stagnation point moves as it does.
The rear stagnation point is usually located at the trailing edge of the airfoil, thanks to the Kutta Condition which says that, "A body with a sharp trailing edge which is moving through a fluid will create about itself a circulation of sufficient strength to hold the rear stagnation point at the trailing edge." (according to wikipedia). The airflow around the airfoil, which is a body with a sharp trailing edge, keeps the rear stagnation point where it is.
This moving of the front stagnation point as the AoA increases is part of the reason why lift increases with increases in AoA.


How should the number of maintenance hours per flight hour for the C-17 be interpreted? While reading up on the Boeing C-17 Globemaster III, I came across a reference to its maintenance man-hours which surprised me. As listed in the spec sheet (emphasis mine):

  These requirements include an aircraft mission completion success probability rate of 92 percent, only 20 aircraft maintenance man-hours per flying hour, and full and partial mission availability rates of 74.7 and 82.5 percent, respectively.

Now this strikes me as a rather steep maintenance requirement, especially considering the spec lists "only 20 hours". If I understand that statement correctly, for every hour spent in flight, 20 technicians have to spend that same amount of time maintaining the plane after landing.

Do those 20 hours refer to an average over extended operation of the aircraft? In other words, does this include extended maintenance where the plane might be undergoing days of intensive maintenance, skewing the numbers?
Is this a low number for this class of aircraft?
Do the required maintenance hours scale with: size, complexity, military/commercial use?

I am sure the raw data can be found somewhere in some document or budget line item but there are a few things to keep in mind.

These numbers may only be airframe maintenance preformed. According to this article it seems that P&W maintains the engines.
This article (a bit dated) states that the air force must only maintain 50% of the work in house. Again they may only be counting their hours.
The air force (and all military aircraft) are subject to their own regulations and intervals and such. Although they must play by FAA rules when flying in airspace their upkeep is to the standards set forth by the owning military body. This may make it hard to compare with general aviation or commercial aviation.

  Do the required maintenance hours scale with: size, complexity, military/commercial use?

In general yes. A piper cub is going to be simpler to maintain per hour than say an F-16 but as you know things vary. I would say that the hours scale more with complexity than size. Mainly because from a mechanical stand point a more complex engine or aircraft would have more things to disassemble and go wrong and in turn more things to put back together generally increasing maintenance time. Keep in mind this is a very general statement.

In terms of military vs commercial use, you will see a lot more maintenance on a military fighter than a commercial plane largely as a result of design. Commercial planes have lots of redundant systems and reliable parts as they are selling safety. Fighters have lots of high performance parts, some redundancy and things that may be made to wear faster in order to preform. Again a general statement but you can get the gist.

  Do those 20 hours refer to an average over extended operation of the aircraft? In other words, does this include extended maintenance where the plane might be undergoing days of intensive maintenance, skewing the numbers?

Im going to assume that this is an overall average but it may not be.


What's the point of separating the Alternator and Battery switches? What's the point of having a different button for the battery? Why does the alternator turn on simultaneously with the battery? I need an easy and short explanation, if possible.

I will expand a bit on @Noah Krasser's answer with some situations where you would want one or the other. FWIW under-volt scenarios are typically far less of a problem than over-voltage scenarios.

Bad Voltage Regulator: Alternators vary their voltage output with RPM. To keep your system operating properly a voltage regulator is installed on the alternators direct output. This keeps your systems power supply at a constant voltage, typically at the cost of some power in the lower end. These things can and sometimes do fail, if they fail in such a way that the alternator is at a high RPM and putting out a voltage that could be potentially harmful to your avionics you are going to want the ability to cut it off while still drawing from your battery.
Battery Short: More often than not batteries are stored in the tail section of GA planes or potentially anywhere. If the cable chase leading to the battery develops a short you may want the ability to cut the battery circuit out while still having the ability to power you aircraft from the alternator.
Thermal Runaway: Batteries can runway on you and cause an over-volt issue. Again in flight you are going to want to be able to cut it from the system.


Why do military turbofan engines use a low bypass ratio? I know that the most civilian engines use a high bypass ratio which is good for fuel economy and noise reduction.

What prevents military engines from using the same technology instead of opting for low bypass engines?

It's not military vs civilian, but subsonic vs supersonic-capable

Note that subsonic military aircraft use the same engines as civilian aircraft, even if their names might be different.

The KC-135 used initially the J-57 which was called JT-3C when used in the Boeing 707-120. Now they fly the CFM-56, which is used on the Boeing 737 and the A320.
The C-5 Galaxy uses the GE TF39 which became the CF6 when mounted to a Boeing 747-100 or a DC-10.
The Fairchild A-10 uses the GE TF34 which is called CF34 when mounted to civilian aircraft like the Bombardier Challenger.

No, the differences arise only when the aircraft is designed to fly supersonic. This requires a very different approach to the integration of the engine:

Supersonic aircraft engines are mounted close to the centerline. If possible, they are straight behind the intakes, so the intake flow does not need to change direction. Exceptions like the SR-71 are rare.
Supersonic intakes are longer and have sharp edges as opposed to the short, blunt intakes of subsonic aircraft. Also, most have a variable geometry to adapt to the very different flow conditions at supersonic speed.
Since it is the job of an intake to slow down the air going into the engine, supersonic intakes cannot have a big capture area, or their spill drag in supersonic flight would be excessive. Supersonic engines need to create their thrust with much less airmass than purely subsonic engines. Forget stealth, this is the real reason for the smaller diameters of supersonic-capable engines.
The nozzle of a supersonic aircraft is also variable, in contrast to the fixed nozzle of subsonic aircraft. This again helps to adjust it to the flow conditions, but in this case the major difference is between reheat on and off. Afterburning engines are capable of much higher exit speeds to compensate for their smaller diameter. They accelerate less air to a higher speed to create comparable thrust.
The last point mentioned it, but it deserves a bullet of its own: Supersonic engines use afterburners in order to have enough thrust for going supersonic at all. The hot exhaust gasses have a much bigger volume than the cold intake flow which needs to be accommodated by widening the nozzle.

Note that the civilian Concorde used also a variable intake and nozzle and afterburners. It had an engine which was used on the BAC TSR-2 before, a supersonic military aircraft.

The real distinction is not between civilian and military, but between purely subsonic and supersonic-capable. Initially, both was achieved with the same engines. The J-57 mentioned above was also used on the supersonic F-100 military jet. Only in the 1960s did those lines diverge, and the subsonic aircraft grew ever bigger low-pressure compressor stages. These were again driven by the high-pressure cores which were used on supersonic aircraft.

Background

Thrust is air mass flow multiplied by the speed difference between flight and nozzle speed of the engine. To increase thrust, subsonic engines try to maximize mass flow (by increasing the bypass ratio) while supersonic engines rely more on increasing the nozzle speed (by using afterburners). Since net thrust is only possible when exit speeds are higher than the flight speed, the engine's exit speed needs to increase with the design flight speed.

The core engines do not differ much - after all, the intake will make sure that air reaches the engine at a speed of Mach 0.4 to 0.5, regardless of flight speed. The core of the General Electric F110 (installed in the F-15 and F-16 fighters, among others) became the core of the CFM-56 turbofan which is used in the Boeing 737 or the Airbus A320. The main difference is in their bypass ratio. The slower the design speed, the bigger the bypass ratio may become. At very low speed, the ungeared, shrouded fan is exchanged for a geared, free spinning propeller, in other words, the jet changes to a turboprop. The intake and nozzle, however, are very different indeed.

The optimum bypass ratio changes continuously, but since the drag coefficient drops after crossing Mach 1, airplanes are either designed for a maximum Mach number of 0.9 or less, or 1.6 and above. The corresponding bypass ratios today are up to 12 for subsonic engines, and less than 1 for supersonic engines. This produces a sharp boundary at the speed of sound, and many military engines designed for supersonic flight lost their afterburners and were fitted with a big fan to become the engines for subsonic transport aircraft.

The differences between sub- and supersonic engines grow bigger the more you move away from their core. High-pressure compressor, combustion chamber and high pressure turbine look and work the same, but the low pressure compressor of subsonic engines swallows a lot more air and has a much bigger diameter. Supersonic engines in turn mostly have an afterburner. The biggest difference, however, are the intakes (large pitot intake with blunt lips for subsonic aircraft versus adjustable spike or ramp intakes for supersonic flight) and the nozzle (fixed for subsonic flight versus a complex, adjustable convergent-divergent nozzle for supersonic flight). This is due to the very different air speeds and the much higher exit velocities required for supersonic flight.

Look at the intake section of the XB-70 pictured above (source). The capture area is rather small, and then the intake tube widens to enable the slowing of the airflow. The inclined sidewalls of the intake section cause a lot of drag at Mach 3. Now think the six GE YJ-93s are replaced by engines with an even bigger diameter. The increase in wave drag due to the even blunter intake would cancel out all advantages of a higher bypass ratio.

What is less obvious is the fact that this intake section also creates maybe half of the overall thrust of the propulsion system. But this answer is already too long, so I save this for another answer.


For any given modern turbofan engine, does bypass ratio stay approximately constant during a typical flight? Bypass ratio is the quotient of mass flow rate through the bypass duct over mass flow rate through the engine core. Does it stay constant during the typical flight of an airliner?

If it does not stay virtually constant, what are the driving factors?

There are two factors that determine the bypass ratio (or rather, how it may change):

The first is the relative diameter of the fan, compared to the core. Obviously, this is fixed in the design process, and doesn't change during flight
The second is the relative difference in RPM speed between the two (or even 3) shafts that exist in the engine. This does change during a flight, from the takeoff condition, to a lower thrust setting at cruise. So let's look at this in detail.

The compressor acts like a pump, increasing the pressure of the air, whether it's sucking at the intake like in takeoff, or RAM feed, such as in cruise flight. Also like a pump, the faster a spool (compressor) spins, the larger the mass flow it pumps (creates). So, if we consider a CFM56-7B engine for example, its shaft speeds at 100% are 5,175 rpm for N1 and 14,417 rpm for N2. That's a ratio of 2.78. At takeoff, the engine will be close to this condition. But at cruise thrust at 35,000ft, the engine will be throttled back. What tends to happen in a gas turbine, is the N1 spool slows down relatively more than N2, as the throttle is reduced. So at cruise, N1 is maybe 3,880 rpm (75%), while N2 is maybe 11,534 rpm (80%). The ratio between them is now 3.41. The pumping capacity of the core has only reduced 20% from the takeoff condition, while the pumping capacity of the fan has reduced 25%. Hence, the bypass ratio will be lower at cruise than at takeoff.

Now, my values of 80% and 75% at cruise are just suggestions. But it is to illustrate the point, that if the ratio between the rpm of the two spools change, the bypass ratio will change.

In essence, the core flow is being determined by the rpm of both spools, while the fan duct flow is just determined by the rpm of the fan. So changes in the rpm ratio will impact the bypass ratio.

{Why is the N2 operating range less than the N1 range?: The CFM56 engine has a single stage fan, and depending on the model, either a 3 or 4 stage booster (or LPC), with a 9 stage HPC. These compressors are making the air flow in the direction of increasing pressure. The air doesn't really want to do that. So, if the design tries to increase the pressure ratio per stage too much, the blade is likely to stall. Getting aerodynamic conditions which avoid stall for the first stage, and the last stage, for the same operating condition is difficult. It's even harder when there are 9 stages on the shaft, than 3 or 4 (plus a fan). So the range of conditions that a 9 stage compressor will operate over is generally less than the range of conditions that a 3 or 4 stage compressor will operate over.}


Boeing 787 type rating How tough/easy is to crack a 787 type rating course? Is a wide body aircraft difficult to fly especially when you are transitioning from 737? Any tips/suggestions would be much appreciated.

Biggest things here are going to be the availability of training courses as well as full motion simulators to do the training in. There are a lot of FMS systems for 737 aircraft operated by multiple companies offering type ratings in that airplane.  Conversely I only know of a few companies that offer initial and recurrent type rating work in a 787.   Boeing offers a type rating course in this aircraft, and as well as the individual 787 operators have their own initial and recurrent type rating training.

Your ease of obtaining a type rating is largely going to be a function of how big your bank account is.  You will also have to select whether you want to receive an initial type rating as a first officer on 787 or as a captain of the 787.  The training will probably take 3-4 weeks and approximately 20 to 30 hours of sim time (you will never fly the real thing for a type rating) to complete in a full motion simulator, and those are costing roughly 1000 dollars an hour to operate.   So at that time requirement plus say 2000 dollars for the course cost of materials,  you're probably looking at around 22,000-32,000 USD, provided of course you pass the check ride at the end.   If you were new to jets or other fast movers, this would probably be a monumental task for you.  If you have been type rated already in another airplane like the 737 or similar, the process should not be very difficult.

On top of that, unless you're going to go immediately fly for a 787 operator somewhere, that type rating is pretty much useless. (If you don't already have 4000 hours TT plus at least 1500 PIC Jet time, don't even bother applying; no airline would hire you for that role, type rated on the jet or not).  And it's only good for 12 months; you'll then be required to undergo a recurrent type rating as either a first officer captain on the 787 and will cost about 60% what the initial type rating did.


What does the zig-zag pattern on Hawker Hunter's leading edge represent? On some Hawker Hunters, there is a zig-zag on the leading edge of the wing, as shown below.

Why do only some Hawker Hunters have this feature, and what is it for?

After doing a little more research I found the answer.

The zig-zag on the leading edge is a type of leading edge extension called a dogtooth extension.

Quote from wikipedia:

  A dogtooth is a small, sharp zig-zag break in the leading edge of a
  wing. It is usually used on a swept wing, but also on straight wings
  ("Drooped Leading Edge" arrangement), to generate a vortex flow field
  to prevent separated flow from progressing outboard at high angle of
  attack. The effect is the same as a wing fence.

  Where the dogtooth is added as an afterthought, as for example with
  the Hawker Hunter and some variants of the Quest Kodiak, the dogtooth
  is created by adding an extension to the outer section only of the
  leading edge.


What would be the expected approximate ROD when descending at 250 kt IAS using LVL CHG at 10,000 ft, assuming a GW of 60t? The question refers to a B737NG but I suppose most similarly sized airliners would show a similar ROD under the circumstances.

A descent at idle power and 250 knots yields about 1,500 feet per minute in the 737. Perhaps a bit less in the Max (maybe 1,300 - 1,400 fpm) with its lower drag profile, and a bit more (like 1,600 fpm) in a Classic.
Source: personal experience flying 737's.


What are the EASA external aircraft lighting requirements for day/night/VFR/IFR operations? I have attempted to decipher the EASA rules related to non-commercial small aircraft external lighting requirements (for a pilot) for the various types of flight, and I can only find the MEL for complex aircraft.  Can someone point me to the requirements?

For EASA (via europa.eu, Easy access version), which is very similar to ICAO Rules of the Air (Annex 2):

SERA.3215   Lights to be displayed by aircraft
(a) Except as provided by (e), at night all aircraft in flight shall
display:

anti-collision lights intended to attract attention to the
aircraft; and

navigation lights intended to indicate the relative path of the
aircraft to an observer and other lights shall not be displayed if
they are likely to be mistaken for these lights; or

in the case of balloons, position lights.

(b) Except as provided by (e), at night:

all aircraft moving on the movement area of an aerodrome shall
display navigation lights intended to indicate the relative path of
the aircraft to an observer and other lights shall not be displayed if
they are likely to be mistaken for these lights;

unless stationary and otherwise adequately illuminated, all
aircraft on the movement area of an aerodrome shall display lights
intended to indicate the extremities of their structure, as far as
practicable;

all aircraft taxiing or being towed on the movement area of an
aerodrome shall display lights intended to attract attention to the
aircraft; and

all aircraft on the movement area of an aerodrome whose engines
are running shall display lights which indicate that fact.

(c) Except as provided by (e), all aircraft in flight and fitted with
anti-collision lights to meet the requirement of (a)(1) shall display
such lights also during day.
(d) Except as provided by (e), all aircraft:

taxiing or being towed on the movement area of an aerodrome and
fitted with anti-collision lights, to meet the requirement of (b)(3);
or

on the movement area of an aerodrome and fitted with lights to
meet the requirement of (b)(4);   shall display such lights also
during day.

(e) A pilot shall be permitted to switch off or reduce the intensity
of any flashing lights fitted to meet the requirements of (a), (b),
(c) and (d) if they do or are likely to:

adversely affect the satisfactory performance of duties; or

subject an outside observer to harmful dazzle.

You might want to refer to a country's specific AIP just to be sure that there aren't local differences.
Not to be used in lieu of the above, AIP, or airport textual data, what follows is my summary by light type:

navigation lights:

in-flight at night
moving on movement area at night
stationary and inadequate area illumination

anti-collision:

in-flight by day and night
taxiing; being towed -- unless too dazzling; generally in busy controlled airports the ATC will ask for them to be switched off; see (e)

red beacon (or similar):

engine running day or night on movement area


How often can I fly my own small plane? I was wondering about how often can one fly a privately owned plane? Is it somehow like owning a car where you just light up the ignition and just drive away? Obviously, there are a lot of regulations and rules... but say I have my plane parked in a hangar in a nearby airport. Can I just drive up to that airport and take my plane for a spin anytime I want (after coordinating with the control tower)?

Depending on what country you live in, yes, but it will cost you.

It's probably the cheapest to fly an airplane you own in the United States.  Here, the answer is yes, with very little restriction, you can just drive out to the airport and fly an aircraft within the National Airspace System.  You just must abide by the legal and medical requirements for you, and the legal and operational limitations of the aircraft.  Most airports throughout the country are uncontrolled, so you may have limited, if any, interactions with ATC.  The ability to fly pretty much anywhere you like with very few restrictions is what makes GA so appealing in the United States.


Is a propeller a wing? An aviation expert claimed to me that a propeller is a wing.

While I understand that propellers use similar principles to generate force, it muddied the definition of a rotary wing aircraft vs a fixed wing aircraft.  If the propeller of a fixed wing aircraft is considered a wing, what is the distinction between the two?  Is it how much surface area each type of wing works on?  Is it the direction of force?

I think most people would say the answer is no. Consider the dictionary.com definition as it relates to Aeronautics:

  9. Aeronautics .
  a. one of a pair of airfoils attached transversely to the fuselage of an aircraft and providing lift.
  b. both airfoils, taken collectively.

The wings are the big airfoils which are responsible for generating most of the vertical lift.

A colloquial test of this definition would be to grab a CFI and ask him/her to "point out the wings on this airplane to me." They won't point out the propeller. (Not a scientific experiment)

However, both wings and propellers are airfoils (and so are the other wing-like surfaces such as the vertical and horizontal stabilizers). You might get some disagreement about whether those other surfaces are wings, but I believe the answer is still no. There is nothing to be gained by calling them wings, and only leads to ambiguity. We already have a word which describes lifting surfaces in-general, and again, it's airfoil.


What is preferred in an emergency landing: a road or a cornfield? When faced with the situation of a power off emergency landing, how do you best select your location to land?

At what point is a road too small or the crosswind too great to attempt a safe landing and instead select a corn field or such? Should the road even be considered?

EDIT: Also, how does plane size and build factor into the decision?

Really depends on what you are flying, what's growing, what season and what the roads are like. Our Cessnas have used gravel roads a couple of times, either due to problems aloft or at the airport (bad weather). In the Alberta prairies finding a side road to land on is no problem whatsoever. 100 and 200-series Cessnas have fairly narrow fixed gear that can take an amazing amount of abuse, don't know about other types.

A real cornfield, with the corn fully grown, will destroy a light aircraft. The plants are 2-3m high with a hard, heavy log of corn at the top. Hit a couple of thousand of them at 100km/h..... (Skydiving centers in corn country have firm advice for landing in the corn: keep your legs crossed.)

Ripe canola is also nearly 2m high and is a bushy plant interwoven with it's neighbors. You'll probably flip over forward. Wheat is the best choice.


What kind of cable is used to transmit signal from flight control computers to actuators? I am a student and have been researching about ARINC 629 and AFDX. In b777 fly-by-wire system, what kind of cable is used to transmit analog signal from Actuator Control Electronic (ACE) to the respective electrically-controlled actuators? What kind of bus cables/connectors are used in ARINC 629?  What about in AFDX? Does AFDX use fiber optics?

ARINC 629

From The Avionics Handbook, chapter 11, Boeing B-777, Fly-By-
Wire Flight Controls:

  The ACEs [Actuator Control Electronics] and PFCs [Primary Flight
  Computer] communicate with each other, as well as with all other
  systems on the airplane, via triplex, bi-directional ARINC 629 Flight
  Controls data busses, referred to as L, C, and R. The connection from
  these electronic units to each of the data busses is via a stub cable
  and an ARINC 629 coupler. Each coupler may be removed and replaced
  without disturbing the integrity of the data bus itself.

Triplex, bi-directional ARINC 629 Flight Controls data busses

Triple: The B777 has A629 buses that can be dual, triple (flight controls) or quadruple (engine controllers) for redundancy. A triple bus means the information is send on 3 buses in parallel, using 3 bus interfaces, like this:

Source

The primary flight controllers have three lanes allocated to different functions: command, monitor and standby. Each lane is connected to a different bus of the triplex.

A629 is in many points similar to thin Ethernet (e.g. 10Base2) but uses a twisted pair instead of coaxial cable for the physical layer:

Bi-directional: Contrary to A429, A629 is bi-directional, data can travel in any direction,
From any interface to any other interface,
The bus is a twisted cable,
Terminators are fitted at both sides to prevent signal reflection and stationary waves,
A stub connects the unit to the bus using a coupler,
Couplers are plugged directly on the cable without cutting the wires.

Source

Coupler: This is actually a transformer and data are transmitted by induction.

Source

Data: Data on the bus are digital, coded with Manchester bi-phase.

AFDX

AFDX (Avionics Full-Duplex Switched Ethernet) is similar to 802.3 Switched Ethernet Standard. The physical layer can be a shielded twisted pair or an optical fiber.

Instead of a bus topology it uses a star topology. Switches are used to connect the different nodes.

Source

The particularities of the switched network:

Data don't flow to each units on the same bus, the network can be segmented to reduce contention (collisions).
Since the broadcast domain is reduced, this allows for higher data rates.
Deterministic transfer times can be achieved, something which is important in real time applications.

AFDX also allows for virtual links, that is defining a point to point virtual circuit, to completely abstract the network layer.


Starting Off My Career as a Pilot I am 17 years old and am interesting in pursuing a career as a pilot. I am not sure what I need to do first, but I do have a flight lesson this Saturday to see if I want to continue with this career plan.
To start off, what certificate should I work toward first and where should I go from there? I have read about the recreational pilot's license, as well as some other licenses, and was wondering what I should go for first.
Also, any links to study material for any exams I may need to take would be greatly appreciated! All of the information about becoming a pilot seems a bit overwhelming, so I am just looking for a place to start and a basic plan.
Thanks!

If you're considering going career, skip Sport and Rec; you'll need Private before anything higher, and there's no credit for having anything lower first. But wait until you see how your first few lessons go before getting too far ahead of yourself.


How is proper CG maintained in ultralight pusher aircraft? Considering pilot weight can vary 100lbs or more how do ultralight pusher airplanes maintain proper CG range?

Usually with suitably-placed counterweights. I used to fly a two-place pusher Chickinox kot-kot that, to be flown alone, needed a lead counterweight of about 15kg fastened on the extreme nose.


Why Consumer Drones don't use multi-blade (more than 2) propellers? Maybe unusual question but it's related to Aviation world, even if is more directly related to Drones one.

Why almost all consumer drones like DJI's and Yuneec's ones (but it's true also for other manufacturers) uses two blade propellers only?

I'd think that tri-blade or quad-blade could spin slower and so quiter.

Exist 3rd party multi-blade propellers compatible with commercial drones, so is possible to use them but what advantages/disadvantages they have? and again... why default ones are only two-blade?

There are a number of related questions on the site for more information about this. I'll link them below.

The short answer is that two blades will be the most efficient use of energy. If the drone needs a better climb rate then adding extra blades is an option, since you can't make them any longer. The downside is that it will reduce efficiency and thus your battery time.

I can't find the question at the moment, but Peter Kämpf mentioned in an answer that theoretically one blade would actually be most efficient except that you would have to provide counterweight to balance it. The counterweight with no excess thrust would negate the advantage, but that's just an illustration of how the fewer blades the more efficiency.

Here are the links for more reading:

What's the difference between a two-blade prop and three-blade?

Is a propeller with two long blades more efficient than one with three shorter blades?

How does blade solidity ratio relate to thrust/power/torque of a propeller?

Why are three-bladed helicopters relatively rare?


Is there any limit to how many times an aircraft evacuation slide may be deployed? I've seen multiple videos online of evacuation slide tests and they all tend to be almost violent as they are very fast and loud, but I can't help thinking that it must put a lot of force on those seams when they inflate and along with that the material. Is there any limit to how many times these may be deployed before they must be retired?

The slide has to be removed from the aircraft completely (generally) and sent back to it's manufacturer to be inspected and repacked, but yes, it can be reused.

This article talks about it in more detail.

As a far as "how many times", I would assume until it breaks ;).  Though honestly, I doubt a slide would ever be deployed more than once, maybe twice in it's life span.


How do aircraft manufacturers determine the elevator size? I’ve heard some pilots say the MD-11 (I think?)was made with too much elevator, and that it’s an overkill.

So now I’m curious, how do they determine the length for the elevator? Is it possible to make one that can be called an “overkill” without it having drastic effects on the flight?

There are many factors going into the sizing of elevators, and they are all interconnected with other considerations such as CG envelope, static margin, flight controls, etc.

Here are some factors that will influence up-sizing the elevators. Usually, the minimum elevator sizing is determined by low speed requirements:

Ability to perform early rotation in both all engine operating and one-engine-inoperative conditions (14 CFR 25.107(e)).
Ability to perform go-around in the highest flap extended position (14 CFR 25.119).
Ability to stall the aircraft to demonstrate the intended stall speed (14 CFR 25.201). This may not be obvious, but if you lack the elevator authority to demonstrate the intended stall speed, then you can only declare what you have demonstrated, which negatively impacts all the performance speeds for takeoff and landing.
If the elevator is power actuated (e.g. hydraulically powered), then you also need to consider regular takeoff, landing, go-around, and maneuvering at low speed, with failures (14 CFR 25.671(c)).

Here are some factors that will influence down-sizing the elevators:

Demonstration of static longitudinal stability at Vmo/Mmo. For column based inceptors, the average stick force to speed change must be no shallower than 1 lb per 6kt (14 CFR 25.173(c)). For FBW aircraft, too large an elevator can also lead to residual oscillations.
General over-sensitivity to control inputs at high airspeed.

Of course, most of these aforementioned points are heavily influenced by flight control design, such as anti-servo tab, bob weights, artificial feel force, etc.


Was destruction of SR-71-related tools really extraordinary? While reading about Lockheed SR-71 on Wikipedia I was baffled by a piece of information regarding Robert McNamara's order to destroy tools used to create Blackbirds, which took place in 1968 (at the beginning of SR-71's service), to prevent them from falling into anyone's hands.

While post-production cleanup of tools by planes manufacturers seems like a normal process (why waste space, resources, taxes, etc.), it appears that creation of any parts was impossible.
Although destruction of tools doesn't sound like a good idea to secure the technology (I somehow doubt the govt gave up on know-how and general knowledge - as for what can happen B-2's history can be a notable example), I'm still wondering:

Were all the tools indeed destroyed? If so, how were the planes maintained for all those years?
Were in said order any remarks regarding destruction of knowledge?
Why weren't the tools reconstructed after a few years?
That order sounds like a really bad idea - was there any legitimate reason for it?

was there any legitimate reason for it?

Unused tools, the tools for making a plane which isn't made anymore, are most probably put somewhere in storage, since the factory floor is best used for actual production.

While an aircraft plant for making top-notch, spyplanes is generally quite secure, any storage facility is probably less secured. Maintaining a high-security level for a storage facility seems like spending good money for a not awfully good purpose.

So in order to prevent the tools for the secret planes being stolen from a run-of-the-mill storage facility, it is much more secure to have them dismantled.

Note, that the SR-71 was produced in a time, where 24/7 surveilance was not as common as it is today. I suspect, that setting up a high-security storage facility would involve a lot of guards guarding basically junk-metal. Guarding scrap metal also doesn't seem to be an awfully motivating task for the guards.


How much can falling down due to a stall influence the angle of attack? I am trying to understand the af447 incident, more precisely, how a professional pilot having an attitude indicator available failed to understand that the angle of attack was too high. Was it that the angle of attack may be too high at the horizontal attitude, or close to that? As I understand, the aircraft stalled due to pilot error.

Over the longer range it cannot be just a free fall; the air friction would stabilize even a round ball to some stable falling velocity.

Knowing the downward velocity, as well as forward velocity, we can compute the angle of attack for horizontal attitude, and various different attitudes. Seems that if it falls with the same speed as it moves forward, the angle of attack would be 45 degrees just for a normal horizontal orientation. Is this situation realistic?

In AF447 the falling motion most definitely affected the angle of attack. Due to the fact that the pf did keep the stick pulled back for all but the last few seconds of the fall the aircraft was so steep that, although the aircraft was falling at a slight pitch up, angles of attack actually reached as high as 60°.

The descent was so steep that the air was passing over the pitot tubes at such an angle that it was reading below 40 knots, although the plane was moving far faster then that. The computer required that the airspeed be of a certain value to consider the aoa plausible so the stall warning stopped sounding even though the aoa was between 30° and 60° during the fall. When it hit the water the forward ground speed and the vertical speed were both 107 kts.

As far as why the pilot did not realize he was stalled, that's the question of the century. Much effort has been put into figuring that out, and there are numerous things that are at play. But in the end I don't think anyone has come up with a satisfactory answer for that.

If you want to dig into the causes of this crash and get a good understanding of Airbus systems in general I highly recommend this book and it's companion website. The author is uniquely qualified and gives the most thorough explanation yet is not as technically challenging as reading the BEA report.

All of the numbers mentioned in this answer come from the linked book.


What are the Ikarus C42 wing loading limits? I have been reading about the Ikarus C42 aircraft as a potential personal recreation aircraft.  It compares very favorably with the Cessna 152, most notably in weight, stall speed, and fuel economy.

Does anyone have info on the negative and positive G loading limits for this aircraft? Can it be dis-assembled for transportation and storage? Thanks.

My club has a C42B that is used both as a trainer and a recreational aircraft, so I stopped by and looked in the Pilot's Operating Handbook for that one.

Now, before I say anything else: obviously, if the exact values matter, then always consult the POH for the specific airframe in question!

With the above preliminaries thus out of the way:

The structural limits are stated as maximum positive wing loading +4 G, and maximum negative wing loading -2 G.
The wings can be removed, and the POH lays out the procedure for this (as well as, obviously, that for reattaching them).

At least while quickly glancing through the text, I didn't spot anything obvious which would indicate that this differs between variants in the C42 series.


What does Airspace 24 mean in this chart supplement? I found this line in Leesburg Executive Airport (KJYO)'s chart supplement:

I have looked in the Chart Supplement Legend under section 29 and there is no information about such an airspace. At the bottom of the legend it says to look at AIM's Chapter 3 for more details, but none can be found there. I googled this to no avail.
I am guessing that whatever it is might have something to do with KJYO's situation. It is located in the DC SFRA and has an FDC NOTAM "Leesburg Maneuvering Area" which lists required procedures when operating in and out of the airport. Don't know how "24" comes from all that, though.

It looks like a mistake of some kind. The next release of the Chart Supplement is valid from 2021-04-22 and the airspace line has been removed completely from the KJYO entry.


Why does the aspect ratio of a wing become less important at supersonic speeds? My understanding is that an increase in the aspect ratio (AR) of a supersonic wing won't increase the efficiency like it would for a subsonic wing.

If the subsonic induced drag is:
$C_{D_i} = (C_L^2)/(\pi\times AR\times \epsilon)$

And the supersonic lift wave drag for slender bodies* is:
$C_{DWL} = 2\times \frac{C_L^2}{\pi\times AR}$

They're both proportional to $\frac{1}{AR}$ so why is the aspect ratio considered insignificant at supersonic speeds? ------> It seems like I'm missing something here.

*Does this term only apply things like fuselages and nacelles and not wings? If so, what does it mean for it to have an aspect ratio?

Peter Kampf please help

The two main reasons for wings with a subsonic leading edge are:

The lower lift coefficient in supersonic flight, and
the need to keep wing thickness down in order to minimize wave drag in supersonic flight.

Your observation is correct: The lift-related drag coefficient is proportional to the inverse of the aspect ratio, regardless of flight regime for slender bodies. In all cases, lift is created by accelerating air downwards, and efficiency is better when more air can be accelerated by less.

But another factor is the lift coefficient, the square of it to be precise. Again this influence on drag is equal for sub- and supersonic flight, but the magnitude of the lift coefficient isn't. Take the XB-70, for example (source):

Cruise lift coefficient:  0.1 to 0.13, $c_L^2$ = 0.01 to 0.0169
Takeoff lift coefficient: 1.3 to 0.73, $c_L^2$ = 1.69 to 0.533
Landing lift coefficient: 0.626, $c_L^2$ = 0.392

Even at the highest supersonic cruise lift coefficient, the approximated lift-dependent drag coefficient is only 4.3% of what it is at landing speed. Supersonic drag is dominated by pressure and friction, and the cost of lift creation is dwarfed by the other sources of drag. Or let's look at another aircraft designed for sustained supersonic flight, or more specific, it's flight envelope (source):

The aircraft operates in a narrow band of dynamic pressures, trading density for speed to stay within a corridor of 310 KIAS to 450 KIAS when flying at supersonic speed. Here the lift coefficient is roughly between 0.1 and 0.3 (depending on aircraft mass and load factor) and again much lower than at low speed, where minimum speed drops to 145 KIAS.

Note that the lift wave drag formula for slender bodies is only strictly true for $AR$ approaching zero. At higher aspect ratios and especially with a supersonic leading edge this relationship is diminished. In supersonic flow the aspect ratio only comes into play by influencing the part of the wing that is within the Mach cone of the wingtip. In other words: The supersonic flow around a slender body is everywhere influenced by tip effects, whereas the inner and forward part of a higher aspect ratio wing experiences two-dimensional flow, being blissfully unaware of lateral limitations to wingspan.

Now to the structural factor: Wave drag has two components, one from lift and the other from the displacement effect. Pushing air out of the way at supersonic speed must be avoided as much as possible, because it is doubly punished: Surfaces facing in flight direction experience higher pressure than ambient, and those facing backwards experience lower pressure (i.e. suction). In all, zero-lift wave drag increases with the square of the thickness ratio, so there is no space in supersonic wings for the beefy spar a high aspect ratio wing would require.


Can a surrendered type certificate be reacquired by the manufacturer or by another entity, or is it considered to be permanently lost? On occasion, the holder of an aircraft type certificate or supplemental type certificate chooses to surrender it to its local civil aviation authority (for instance, the FAA or EASA), which has the effect of grounding all aircraft covered by the type certificate; can the former holder, or another person or entity, later reacquire the surrendered certificate, or is a surrendered type certificate permanently voided?

From the FAA point of view, the TC is not lost but cannot be reissued. You have to get a new one if you want to become the new TC owner. Order 8110.120 states:

"The surrender of a certificate is a final action. Once surrendered, the action cannot be reversed. A certificate cannot be reissued to a third party or a former holder. A new TC for the product could only be issued as part of the full TC process with a new application, subject to all current regulations."

I don't know EASA's policy but I suspect it likely mirrors the FAA's.

In addition, your statement that aircraft covered by the surrendered TC are grounded is incorrect. Surrendering a TC does not affect the airworthiness status of aircraft produced under the TC. From the same document linked above:

"The surrender of a TC does not affect the airworthiness certificates of existing aircraft. The aircraft remain subject to future AD actions for safety of flight issues that could affect their airworthiness certificates."


How are attitude indicators kept accurate? The attitude indicator is very important for any sort of instrument flying, so it must be accurate at all times regardless of the plane's movements.

I presume that the indicator would be connnected to a gyroscope. However, gyroscopes drift, and a drifting attitude indicator would be pretty useless. Of course, you might have a reset button, but safely using it would require a visual reference to a horizon, which obviously you don't have during IFR flight.

Another idea might be to periodically reset the indicator based on gravity --- many model aircraft use this something like this approach to drive an attitude calculation; however, during coordinated turns, the gravity vector is indistinguishable from the vector during level flight. If a plane continually turns in a coordinated way, such an attitude indicator would be confused in exactly the same way that a human would be, resulting in bad things like the autopilot doing a graveyard spiral. A gravity sensor + accelerometer wouldn't be able to detect the difference between a gradual turn into a steeper and steeper coordinated spiral and simply a gradual gyroscope drift while remaining in level flight.

How are attitude indicators kept from drifting?

A gyroscopic AI has an erection mechanism, which continuously corrects the AI to be upright based on the local level, or downward acceleration vector. The correction rate is generally 3-5 degrees per minute.

The way that the AI corrects itself is a system of pendulous vanes. When the gyro is not upright relative to the local level, centrifugal force pushes open vanes on the gyro's case. Air escapes through the uncovered holes, applying a force to the case, and correcting for precession.

(FAA AC 65-15A)

The second half is somewhat counter-intuitive, but flying in a coordinated turn will not continuously increase the error in the AI. For the first half of the turn, the AI will add error; in the second half, it removes the error. After a 360 degree turn, the precession error will be removed completely. For more detail, see another answer: If I fly a balanced turn long enough, will the AI show I'm level?


Effect of touchdown on angle of attack, tailwheel vs tricycle I have a theory but I can't find any documented confirmation. Please let me know if my theory seems right:
My theory:
During a regular touchdown, a tricycle gear equipped airplane, since the main wheels are behind the center of gravity, rotates around its transversal axis (nose down), thus reducing the angle of attack. This should reduce the plane tendency to rebound (unless the nosewheel touches the ground).
On a tail-wheel airplane, the effect in reversed, on touchdown the airplane rotates around in transversal axis, thus increasing the angle of attack.
Does that make sense?

Yes, it does make sense.
However, the angle of attack increase of the tailwheel airplane would only be possible if it touches down on the main wheels while there still is much clearance of the tail wheel or skid. Normally, the ground attitude of tailwheel airplanes should be very close to their landing attitude, so all three wheels touch down almost simultaneously.


Which are the main aluminium alloys used in airliner construction? The proportion of composites used in airliner construction has increased in the last decades, with the B787 and the A350 being the first ones with composites making up the majority of wing and fuselage structure. From Wikipedia:

  The 787 is the first major commercial airplane to have a composite fuselage, composite wings, and use composites in most other airframe components.

And

  The A350 XWB airframe will be made out of 53% composites: carbon fibre reinforced plastic for the outer and centre wing box (covers, stringers, spars), fuselage (skin, frame, keel beam, and rear fuselage) and the empennage (horizontal and vertical tailplanes).

The vast majority of aircraft in operation (A320, A330, B737, B777) still use aluminium as the main load bearing construction material though. Which aluminium alloys were used for the wings and fuselage of these aircraft?

2000s, 6000s, and 7000s series alloy are very common for aircraft construction over the past 8 decades of metal aircraft design.  Boeing aircraft are largely manufactured from 2000s series aluminum alloy.


Why is aircraft marshalling still necessary? I've seen in a documentary that many airports have people guiding the airplanes to the parking spots (marshals). I've also noticed that the key point is to reach a precise spot on which the nose landing gear has to be placed. Couldn't that be solved with a camera installed on the landing gear itself so the pilots can check whether they go along the lines and finally reach the precise spot?

There are automated systems that provide the same function as a marshal. They are in place at some airports but not all since upgrades cost money and not all airports may see a use for it.

At smaller fields marshals often serve multiple roles aside from just guiding aircraft they may also handle baggage, pump fuel, drive tugs and all the other various ramp jobs. In this case it may be more cost effective to keep a marshal around than instal a costly automated system that only replaces one of their jobs.


What could cause instant nausea when entering larger passenger planes? This may not be the right forum for this question, but I almost always experience temporary nausea when entering the cabin of a larger passenger plane (e.g. A320). It hits me as soon as I enter the cabin and can be quite debilitating. Strangely enough it tends to dissipate almost completely after about 5 minutes inside the cabin.

Does anyone know what might be causing this? My only thought is that is has to do with the air inside the cabin as even the smell of the cabin when boarding (but still on the boarding ramp) causes the nausea to occur.

Additionally, I'm open to any suggestions for how to reduce this effect... I fly quite often and have learned to put it up with this sensation but I would greatly appreciate not having to go through this anymore.

There are a few things that could be causing it but chances are its not the aircraft. When on the ground, with the engines off, aircraft simply have exterior air circulating through them. In the hot weather they are often connected to an external AC unit similarly in the cold weather they are connected to a heater. Cabin smells tend to vary widely and I cant find any data other than anecdotal evidence about how they are controlled and if they are quantified.

Once the engines have started they can use bleed air but if your symptoms are calm after 5 minutes its likely they are still using ground fed air at that point.

You may be experiencing light symptoms of Claustrophobia upon knowing you are entering a small, enclosed space. Some evidence points to the fact this can be cured with repeated exposure so the more you fly you may notice the effects reduced.


What are the advantages of a Spinning Wing (Magnus Effect) and why haven't any been commercially produced? What were the advantages and why was it never produced?

Some earlier prototypes of spinning-wing aircraft were produced but none were successful. The design has some very serious disadvantages. One of the main issues appears to be undesirable gyroscopic effects.

Production

  ... why it was never produced?

Several full-scale prototype aircraft with rotating Magnus-effect wings seem to have been produced:

921-V - Plymouth. 1930.
X772N - Union Aircraft Co. 1931.

It may be that 921-V is the only one to have flown, crashing after one flight.

   The Flettner airplane

  Built in 1930 (USA), the 921-V is reported to have been flown at least once - ending it's short career with a crash landing. Three cylinders with disks performing as winglets driven by a separate engine. Information on this design needed! It's probably the only aircraft equipped with cylinder wings which made it into the air.

From Pilotfriend.com

In practice the effect can be less efficient than conventional alternatives

  In the early 1920's the force from a rotating cylinder was used to power a sailing ship. The idea, proposed by Anton Flettner of Germany, was to replace the mast and cloth sails with a large cylinder rotated by an engine below deck. The idea worked, but the propulsion force generated was less than the motor would have generated if it had been connected to a standard marine propeller!

From NASA

Advantages

  What were the advantages?

The article referenced suggests

  More wing lift and less drag are the major aims of aviation’s researchers. Maybe the Magnus Wing will supply the answers ... The ship is powered by a conventional gas-turbine jet engine, while the drums are revolved by a separate piston engine. Small wings, heavy loads and quick take-offs would be its great advantages.

Disadvantages

The developer of a working model wrote about a few of the disadvantages (try a google translation of that page for more):

If the cylinder rotation accidentally becomes slow or stops, its lift disappears completely. This plane will never be able to glide.
If a gust blows from backside during a slow flight (takeoff or landing), the cylinder wings generate a down force.
The spinning cylinder wings generate a strong gyro effect, which makes it difficult for the plane to change its attitude.

See video and comments.

NASA did some experiments with rotating-cylinders for flaps (not main lift source)

They concluded

  These experiments demonstrated on the one hand the
  effectiveness of such a high lift system but on the other hand the weakness in handling qualities due to the gyroscopic forces in such an aircraft configuration.

From A review of the Magnus effect in aeronautics

The US Army also carried out a study of the use of the Magnus effect in aircraft
VIEW AND PRELIMINARY EVALUATION OF LIFTING HORIZONTAL-AXIS ROTATING-WING AERONAUTICAL SYSTEMS

It is difficult to pick out a simple conclusion as the study looked at a wide variety of systems. For rotating cylinders in wing (RCIW) they wrote

  Such systems do not appear to have merit as STOL devices.


How is the induced drag calculated for a wing with elliptical planform? How is the induced drag calculated for a wing with elliptical planform ? Is this wing shape the most efficient ?

Induced drag is caused by the downward deflection of the air streaming around the wing. The resulting aerodynamic force is tilted backwards by half the deflection angle, and the air flows off the wing with an added vertical speed component, producing downwash. Increasing the downwash angle means increasing both lift and the backward tilt, so the induced drag goes up with the square of the lift produced.

If you want to minimize induced drag for a given lift, this quadratic dependence means the optimum is reached when the downwash angle is constant over span.

  How is the induced drag calculated for a wing with elliptical planform?

The elliptical, untwisted wing has the same angle of attack and the same lift coefficient over span, and produces the desired constant downwash angle.  To simplify things, let's assume the wing is just acting on the air with the density $\rho$ flowing with the speed $v$ through a circle with a diameter equal to the span $b$ of the wing. If we just look at this stream tube, the mass flow is $$\frac{dm}{dt} = \frac{b^2}{4}\cdot\pi\cdot\rho\cdot v$$

Lift $L$ is then the impulse change which is caused by the wing. With the downward air speed $v_z$ imparted by the wing, lift is: $$L = \frac{b^2}{4}\cdot\pi\cdot\rho\cdot v\cdot v_z = S\cdot c_L\cdot\frac{v^2}{2}\cdot\rho$$

$S$ is the wing area and $c_L$ the overall lift coefficient. If we now solve for the vertical air speed, we get $$v_z = \frac{S\cdot c_L\cdot\frac{v^2}{2}\cdot\rho}{\frac{b^2}{4}\cdot\pi\cdot\rho\cdot v} = \frac{2\cdot c_L\cdot v}{\pi\cdot AR}$$
with $AR = \frac{b^2}{S}$ the aspect ratio of the wing. Now we can divide the vertical speed by the air speed to calculate the angle by which the air has been deflected by the wing. Let's call it $\alpha_w$: $$\alpha_w = arctan\left(\frac{v_z}{v}\right) = arctan \left(\frac{2\cdot c_L}{\pi\cdot AR}\right)$$

The deflection happens gradually along the wing chord, so the mean local flow angle along the chord is just $\alpha_w / 2$. Lift acts perpendicularly to this local flow, thus is tilted backwards by $\alpha_w / 2$. In coefficients, lift is $c_L$, and the backwards component is $\alpha_w / 2 \cdot c_L$. Let's call this component $c_{Di}$: $$c_{Di} = arctan \left(\frac{c_L}{\pi\cdot AR}\right)\cdot c_L$$

For small $\alpha_w$s the arcus tangens can be neglected, and we get this familiar looking equation for the backwards-pointing component of the reaction force: $$c_{Di} = \frac{c_L^2}{\pi\cdot AR}$$

If the circulation over span has an elliptic distribution, the local change in circulation times the local amount of circulation is constant, and the induced drag $c_{Di}$ is at it's minimum. If this would be different, a higher local $v_z$ causes a quadratic increase in local induced drag, so the whole wing will create it's lift less efficiently.

  Is this wing shape the most efficient ?

Only if you ask an aerodynamicist would the answer be yes. An elliptic wing will give you the best ratio of lift to drag, which clearly is one way to express efficiency.

In reality, the wing has to lift itself plus a payload, but only lifting the payload should be considered when formulating efficiency. Therefore, pure lift/drag optimization is too narrow. What should count is the best ratio of lift minus wing weight relative to drag. R. T. Jones wrote a NACA Technical Note back in 1950 in which he looked at this problem analytically. Wing weight goes up when much lift is created near the tips, because this lift will cause a disproportional root bending moment, and the wing spar which has to carry this bending moment is a significant part of the wing structure. Therefore, reducing lift at the tips and adding more lift at the root will create a lighter wing for a modest drag increase, resulting in an overall optimum for an almost triangular lift distribution. When compared to an elliptical wing planform, the total wing span of such an optimized wing is bigger for the same overall drag, but this wing will weigh less.

Spanwise loading comparison for wings of the same fixed lift, from NACA Technical Note 2249.

But this is too easy. Scaling laws must be considered in addition. You know that elephants have much more massive legs relative to their body size than antelopes (or even ants, for an even more drastic comparison), since body mass scales with the cube of linear dimension while structural strength scales only with the square of linear dimension. This means that wing spar weight will be proportionally higher for larger aircraft.

As a consequence, insects have more elliptic wings than albatrosses, and model aircraft have optimum wings which are much more elliptic than the optimum wing of an airliner. The optimum shifts from an elliptic load distribution at very small scales to an almost triangular distribution at large scales.


Why publish an unusable airway? If an airway is marked unusable, why publish it?

Secondary question, can RNAV still be used?

Short answer: only part of the airway is unusable, and even then it's usable with GPS.

Your image shows V522 between FAILS intersection and the ERI VOR but you have to look at the full airway, which runs from the DJB VOR in the south-west to MYPAL intersection in the north-east:

As you can see, only some parts of it are marked unusable, even if they're very large parts in this case. But you can still use the rest of V522 apart from those sections; the NOTAMs say the ERI VOR's signal is unusable beyond 20nm until December:

  FDC 6/6576 (A0050/16) - OH..ROUTE ZOB. V522 FAILS, OH TO ERIE (ERI)
  VORTAC, PA NA ; ERIE (ERI) VOR R-264 UNUSABLE BEYOND 20NM. 19 MAY
  11:59 2016 UNTIL 29 DEC 11:59 2016 ESTIMATED. CREATED:19 MAY 11:59
  2016

A 7-month 'outage' covers several of the FAA's 28 or 56 days chart cycles, so I guess they decided it was worth charting in addition to the NOTAMs. Removing the sections completely from the chart would mean decommissioning or redesigning the full V522 airway, which is presumably a much bigger task than just marking it unusable.

Finally, the ATC orders give some instructions on this and if part of an airway is unusable because of a NAVAID issue they still expect RNAV-equipped aircraft to be able to use it:

  4−4−4. ALTERNATIVE ROUTES

  When any part of an airway or route is
  unusable because of NAVAID status, clear aircraft that are not RNAV
  capable via one of the following alternative routes:

So the only aircraft that can't use those parts of V522 are ones that have to use the ERI VOR signal for navigation and even then they can expect ATC to give them an alternative route.


Why does the local speed of sound decrease with temperature and not pressure? I presume I'm missing something here but from looking at charts indicating the value of Mach 1 at different altitudes it seemed to reduce the higher in the atmosphere you went which I presume is related to reduced temperature and not pressure related which to me seemed counter intuitive. Thanks, hopefully you can clear this up

It may be helpful to remember the ideal gas law here: $PV=Nk_BT$ (the physics version). If we divide both sides by $N$ we have a $V/N$ on the left which is just a $1/\rho$ where $\rho$ is the number density. Thus:

$\frac{P}{\rho}=k_BT$

In an ideal gas, to which air is close enough, the speed of sound is given by:

$c_s = \sqrt{\gamma \frac{P}{\rho}}$

Throwing in our first results, we see that:

$c_s = \sqrt{\gamma \frac{P}{\rho}} = \sqrt{\gamma k_BT}$

In other words, the speed of sound reduces with either pressure or temperature because those two are in turn connected by the ideal gas law. It's basically two different ways of looking at the same question.

Side note: throwing in the effect of density is another question all together since $c_s$ increases with an increase in temperature or pressure but decreases with an increase in density. The question then becomes whether the decrease due to the drop in pressure overcomes the increase due to the drop in density at higher altitude.


How fast can a large passenger jet descend and slow from cruise to emergency land? On 17th April 2018, Southwest 1380, a Boeing 737-700, was cruising about FL320 when it suffered a catastrophic failure to engine 1.  It then took 5 minutes to descend to 10,000 feet, and I noted from the ATC recordings that the pilot said it would require a long approach to land at PHL.

Is this because the airplane could not bleed off airspeed fast enough for a short/straight-in approach? Is this the fastest that a large commercial airplane can descend and slow for an emergency landing?

I assume that, to achieve maximum descent and deceleration the remaining engine would be pulled to idle, and that landing gear would be extended at maximum safe speed, which IIRC is 250 knots?

Would it be reasonable to fly cross-controlled with an aircraft in this condition?  With no engine thrust requiring rudder input, is that safe at all speeds?  How much can it increase drag?

What are the speed limits for extending flaps, and how much do those affect drag?

(I can appreciate that in the case of Southwest 1380 the pilot in command, not knowing the extent of damage, may have been hesitant to do anything that would strain the airframe or controls.  So I'm curious as to how long an "unstressed" descent would take, and then how much faster one could descend by using the most aggressive techniques that a pilot would reasonably use in an uncompromised passenger flight in the same model.)

A 5-minute descent from FL320 to 10,000 feet is pretty quick. As noted in the comments, the descent rate is 4,400 ft/min. For reference, during final approach descend rate is 300~500 ft/min in a small airplane, 500~700 ft/min on an airliner. A typical descent from cruise altitude is performed at 1200~2400 ft/min.
Getting down fast
A "dive bomb" flight-simulator-style descend is not practical if you intend to land. After all, you have to arrest the descend and bring the aircraft to a stable state. You may very well break the plane when trying to pull out of the dive.
Now, what is the fastest way to lose altitude? According to the Boeing 737 QRH, the Emergency Descent checklist (which is conveniently listed on the first few sections) says...

(...) Spreadbrake...FLIGHT DETENT
(!) If structural integrity is in doubt,
limit speed as much as possible and avoid high maneuvering loads.
Set target speed to Mmo/Vmo. (...)

In other words, deploy the speedbrakes, then pitch down to max airspeed.
Note that, however, this procedure is only used when it is absolutely necessary to get down as quickly as possible. The Boeing 737 FCOM also contains the following, in bold letters:

WARNING: Use of speedbrakes at speeds in excess of 320 KIAS could
result in severe vibration, which, in turn, could cause extreme damage
to the horizontal stabilizer.

Landing gear & Flaps
Deploying the landing gear is not good because:

The maximum flying speed with gear extended is 320 knots, while Vmo is 330 knots. Therefore, if you descend with the gear extended, you will lose altitude slower.
The maximum flying speed for extending the landing gear is 270 knots. Therefore, if you wish to extend the gear, you will have to slow down to 270 knots first, then lower the gear handle, wait for the gears to be locked, then pick up speed again.
Flying with the gear extended introduces vibration to the airframe. Deploying speedbrakes already introduces vibration. A damaged engine further adds vibration.
You need the gears to touchdown, so better protect them in the landing gear bay.

Extending the flaps is even worse because the maximum speed for flying with flaps is even lower. The maximum flap extension altitude is also 20,000 ft.
Cross control
Cross-control can probably lose altitude even quicker, but it is not a maneuver you would do on a passenger airliner. When crossing control, the gravity vector is not down, but sideways - meaning passengers will be forced to the side of the seat, a sensation they are not used to. Furthermore, the passengers sitting on the window side will likely perceive the attitude as extreme bank angle and think that the plane is going to crash! The last thing we want in this scenario is a cabin with panic passengers.


What are the applications, advantages and limitations of Gyrodyne rotorcraft? Gyrodyne is a VTOL rotorcraft. It has ability to power both rotors and propellers for generating lift and thrust. What makes it of lesser popularity than helicopters. Are there any applications of Gyrodyne?

https://en.wikipedia.org/wiki/Gyrodyne

The compound helicopter similarly has a separate propulsor for thrust. The gyrodyne differs from it in allowing the lift rotor to autorotate without power during forward flight.

Both are capable of flying faster than a conventional helicopter, while retaining its VTOL abilities. It is thus an attractive idea wherever both high speed and VTOL/hover are required such as military, rescue and executive in-city applications.

However the extra weight and complexity of the thrust system make the already limited, complex and expensive helicopter even more so. Moreover the speed increase over a conventional helicopter is not great enough to be game-changing. Various programmes have led to several prototypes flying, of both gyrodynes and compound helicopters, but there are few if any applications where the high costs and reduced payload justify the modest extra speed over a helicopter.

The Fairey Rotodyne city commute gyrodyne flew but was then cancelled by the UK government.

Several compound helicopters such as the Lockheed Cheyenne have been developed for the US military but then cancelled.

When the USA wanted a fast VTOL military transport, they opted instead for a convertiplane, the Boeing V-22 Osprey. Its speed increase over a conventional helicopter has also turned out to be disappointing.

Maybe one day.


Why did my business class seat have a three point shoulder harness? I recently rode a short haul international flight on Asiana Airlines AAR111 from KIX to ICN. I was surprised that my business class seat had a three-point shoulder harness given that we’ve discussed many times on this SE why shoulder harnesses aren’t provided to customers.

Interestingly, on the flight back I flew Korean on the same route and KAL [business class] seats don’t have three points, only standard seat belts.

Is there a reason AAR111’s business class has three point harnesses? Why just that plane and not similarly configured other flights?

(Source) Your Asiana A350 ride.

On the same plane the economy seats don't have that shoulder strap. It's not new either. There are various reasons:

Most business class seats are not forward facing. For your ride it's a slight angle but it's there.
It allows* reclining the seat during taxi, takeoff, and landing.

  [Qantas will] offer the world’s first seats that let passengers recline in their seat from take-off through to landing. Key to the break-through product’s safety is an over-the-shoulder belt.

It protects the head from hitting the seat walls.
It protects the head since there is no cushion (the knee-breaking seat in front of you in economy).

  (...) the three point seat restraint is used in the larger premium class seat, because there is no seat back close enough to the traveler to cushion his/her head in a sudden stop.

It will keep you in place during turbulence when the seat is reclined.

On the return flight aboard Korean, based on flightaware it was probably a 777-300ER or a 747-8. The missing shoulder strap could be because the seats are forward facing on both types, and/or Korean wasn't interested in the option to recline during takeoff/landing.

* Allows, but not necessarily allowed by the airline, i.e., some airlines may still insist on the upright position for the takeoff/landing.

I also found this supporting text:

  9.20.3 Three-point Seatbelts
  Three-point diagonal fixing seatbelts are occasionally installed in premium seats, including those positioned at an angle away from the forward/aft axis of the aircraft. The intent of the three-point harness is to protect the occupant from contact with the forward side of the seat shell.

— IATA Cabin Operations Safety Best Practices Guide (skybrary.aero).


Is it possible to be a pilot for both the RAF and a Commercial company? I have always wanted to be a pilot for both the RAF and Commercial airlines.
Would it be possible for me to serve, for say 6 months in the RAF, and then when not deployed work as a commercial pilot. Just asking this as I saw a video a few months ago of a US Air Force member who works for United when he is not working for the USAF.
Thanks - Jay.

Basic officer training in the RAF takes 30 weeks, so you're already over your six months, even before you start pilot training.

When you sign up you enlist for a minimum period which varies according to your speciality. Six to twelve years is typical.
Source

Assuming you get through pilot training and make it to a squadron, you're not going to just sit around and wait for a war. There's ongoing training, exercises, active deployments and other duties that will take up your time. Whilst there will be periods where your working life will be 9-to-5, there will also be periods where you'll be working nights, weekends, public holidays, and away from home.

I don't know what the RAF policy on second jobs is, but for practical reasons I can't see that you would be able to drive buses for easyJet or the like.


Do you have to lubricate the shock strut? I have a school report about the lubrication on the shock strut and I was wondering if there is a specific type of lubrication they apply it on.

This is for Cessna 150-152

Oleo struts and brakes used on light aircraft use MIL-H-5606H hydraulic fluid (the red stuff) as the internal damping agent/lubricant (be sure to check the manual anyway). It's a normal practice to clean the exposed chrome of the shock strut (it should be done with the tail held down to extend the strut all the way) with a clean cloth, then wiping it with a cloth with a little bit of 5606 to leave just a very light film to help lube the scraper and seals the next time the strut compresses (once that's done there is the film of internal fluid left behind by the seals).

The biggest enemy of oleo struts is silica particles (which are very hard) from outdoor dust which sticks to the chrome surface because of the fluid film left behind by the internal seals.  The silica collects in the seal cartridge and acts like lapping compound, accelerating wear (this is why desert operations are the toughest environment overall on any airplane).

As the strut ages and the seals wear and slowly start to leave behind more fluid each time the strut extends, dust collects faster and the whole process speeds up, so the strut should be kept clean as much as possible depending on how dusty the local environment is.

If it's your own plane, you should learn to do the published procedure yourself, because nobody else is going to do it between annuals.  If you're renting the plane, that's the owner's problem, and you just do your normal walkaround inspection.