Load Factors, Maneuvering Speeds and Weight

[Landshark]

This is something I have always wondered about since I started training and learned that maneuvering speed went up as weight increased. That was the reverse of what I expected. Since the airplane is heavier, I thought there should be more stress on the wing root or aircraft structure for a given load factor(g). Here are a few limitations for my particular airplane, a 1979 Cessna 152.

Maneuvering speed at 1350 lbs = 93 KIAS
.............................at 1670 lbs = 104 KIAS
Flight load factor (flaps up) ......= +4.4g

My first assumption is that the max load factor of 4.4g is the same at all flying weights.

I will calculate the lift necessary to produce 4.4g at both the lower and max weight.

1350 lbs X 4.4g = 5940 lbs of lift
1670 lbs X 4.4g = 7348 lbs of lift

At a flying weight of 1350 lbs, the wings must produce (and the airframe must withstand) 5940 lbs of lift to generate 4.4g. Any more lift than that would damage something.

But at a flying weight of 1670 lbs, the aircraft structure and wing root have no problem withstanding 7348 lbs of lift.

Why is that so? Where is the flaw in my understanding?

 

[Dan Thomas]

Maneuvering speeds are based on stalling the airplane before exceeding the design G limits. A heavier airplane will stall sooner (its stall speed is higher) and unload the wing. Remember that stall speed rises with increasing load factor; it goes up by the square root of the increase in load factor. A 2G maneuver will increase stall speed by 41%, 1.41 being the square root of 2.

 

[MAKG1]

It actually goes up by the square root of load. That is, load factor * gross weight. For a given weight, it does behave as sqrt(load factor), but we're comparing two different weights.

Take that bit into account, and the numbers make a lot more sense (sqrt(1350/1670) is real close to 93/104, neglecting the difference between CAS and IAS).

 

[Landshark]

When I started to really think about it during the time it took to make the OP, I realized that the maneuvering speed would have to go up with a heavier plane because it is going to take more lift to generate the same load. I guess what I always had a problem with is why the plane can withstand more lift at higher weight.

Since lift is a force, it will be multiplied by the arm to the wingroot, making a certain torque. I don't see why the max limit to that torque increases because the plane is heavier. Or why the airframe can withstand more lift at heavier weights. I am missing something.

BTW, the sqrt of 1350/1670 is a very nice way to think about it. I was gonna go into rhoVsquared and convert the speed and all that to confirm the increase in speed was going to provide a commensurate increase in lift. That was a much more elegant solution.

 

[dtuuri]

Where is the flaw in my understanding?

The limiting value for "lift" is not (+4.4g)*(actual gwt). Instead, it's based on the amount of lift the remaining AoA can suddenly generate. When that "tug" on the lift struts is +4.4g*1670, the max limit is reached. Both weights you have listed will generate the maximum allowable lift at the speeds given and yield +4.4g*1670.

dtuuri

 

[MAKG1]

You're assuming the lift is balancing weight. Generally, when the maneuvering speed is a factor, the aircraft is quite far from straight and level. To get 4.4g, you would basically need to be picking up and dropping the aircraft -- something that might happen effectively in severe or extreme turbulence, or in very strong aerobatic maneuvers. Even a 60 deg steep turn is only 2g.

 

[MauleSkinner]

Since lift is a force, it will be multiplied by the arm to the wingroot, making a certain torque. I don't see why the max limit to that torque increases because the plane is heavier. Or why the airframe can withstand more lift at heavier weights. I am missing something.

The torque on the wing root isn't necessarily the limiting factor...for instance, seats or floorboards may be the limiting factor at lighter weights.

 

[dmspilot]

1350 lbs X 4.4g = 5940 lbs of lift
1670 lbs X 4.4g = 7348 lbs of lift

At a flying weight of 1350 lbs, the wings must produce (and the airframe must withstand) 5940 lbs of lift to generate 4.4g. Any more lift than that would damage something.

But at a flying weight of 1670 lbs, the aircraft structure and wing root have no problem withstanding 7348 lbs of lift.

Why is that so? Where is the flaw in my understanding?

I think what you're wondering is, can an airplane withstand more load-factor (G's of acceleration) at lighter weights if they result in the same amount of lift (wing loading in pounds of force)? The answer is not typically known by pilots and is going to be different for each type of airplane. The criteria the FAA uses in aircraft certification is load factor, so that is what we use and that is how the data is provided to us by the aircraft manufacturer.

Why does the FAA use load factor as the limiting criteria? Well, as someone else wrote, it is not necessarily the load on the wings that determines the aircraft's breaking point. It could be a motor mount, for example. The engine always weighs the same, so 4.4g is the same load on the engine mount regardless of the weight of the rest of the airplane.

 

[Cooter]

I'm not certain what you're asking, but the max load factor will be calculated based on a weight. I would assume it is something close to max gross weight. For fighter aircraft, the G-limit is determined by aircraft weight. So, right after takeoff, you might be limited to 6 Gs, but after some fuel burn, 7.5 will be the limit. Is that what you are asking?

Edit: the maneuvering speed is based on weight and some determined load factor. The lighter you are, the easier it is to get to that G-limit, therefore a slower published speed. That does not mean that is the max capability of the wing, only at what speed the aircraft can get to the max load factor.

 

[Dan Thomas]

The torque on the wing root isn't necessarily the limiting factor...for instance, seats or floorboards may be the limiting factor at lighter weights.

Seats and floors and engine mounts are, by regulation, stronger than the Normal category airframe requirements. They have to survive crashes, not just turbulence or rapid control inputs. See FAR 23: http://www.ecfr.gov/cgi-bin/text-id...e&tpl=/ecfrbrowse/Title14/14cfr23_main_02.tpl

I quote from FAR 23.962:
(1) For the first test, the change in velocity may not be less than 31 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the horizontal plane of the airplane pitched up 60 degrees, with no yaw, relative to the impact vector. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 19g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 15g.
(2) For the second test, the change in velocity may not be less than 42 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the vertical plane of the airplane yawed 10 degrees, with no pitch, relative to the impact vector in a direction that results in the greatest load on the shoulder harness. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 26g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 21g.

Those figures were revised in the 1990s to strengthen the floors and seats, but they were already much stronger than the rest of the airframe before that. 9Gs or something.

See FAR 23.561 for general structural strength required for emergency conditions, whcih would presumably include the engine mount since it's part of the structure: http://www.ecfr.gov/cgi-bin/text-id...5bea73b6&mc=true&node=se14.1.23_1561&rgn=div8


Maneuvering speed is all about stalling the airplane before it breaks, folks. At lower gross weights it doesn't stall as readily under heavy G loads, so Va is lower. See this:

http://flighttraining.aopa.org/maga...ng_Smart_A_New_Look_at_Maneuvering_Speed.html



And for another take on it, which amplifies things a bit regarding certain control regimes: http://www.flyingmag.com/myth-maneuvering-speed

And for the FAA's definition of maneuvering speed, let's look at FAR 23.335, paragraph c:

(c) Design maneuvering speed VA. For VA, the following applies:
(1) VA may not be less than VS√n where—
(i) VS is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum airplane normal force coefficients, CNA; and
(ii) n is the limit maneuvering load factor used in design.


And 23.1507:


The maximum operating maneuvering speed, VO, must be established as an operating limitation. VO is a selected speed that is not greater than VS√n established in §23.335(c).

 

[MauleSkinner]

Unfortunately the additional requirements for seats, floors, and engine mounts don't exist in CAR3, which is what many of our light airplanes were certificated under.

Seats and floors and engine mounts are, by regulation, stronger than the Normal category airframe requirements. They have to survive crashes, not just turbulence or rapid control inputs. See FAR 23: http://www.ecfr.gov/cgi-bin/text-id...e&tpl=/ecfrbrowse/Title14/14cfr23_main_02.tpl

I quote from FAR 23.962:
(1) For the first test, the change in velocity may not be less than 31 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the horizontal plane of the airplane pitched up 60 degrees, with no yaw, relative to the impact vector. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 19g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 15g.
(2) For the second test, the change in velocity may not be less than 42 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the vertical plane of the airplane yawed 10 degrees, with no pitch, relative to the impact vector in a direction that results in the greatest load on the shoulder harness. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 26g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 21g.

Those figures were revised in the 1990s to strengthen the floors and seats, but they were already much stronger than the rest of the airframe before that. 9Gs or something.

See FAR 23.561 for general structural strength required for emergency conditions, whcih would presumably include the engine mount since it's part of the structure: http://www.ecfr.gov/cgi-bin/text-id...5bea73b6&mc=true&node=se14.1.23_1561&rgn=div8


Maneuvering speed is all about stalling the airplane before it breaks, folks. At lower gross weights it doesn't stall as readily under heavy G loads, so Va is lower. See this:

http://flighttraining.aopa.org/maga...ng_Smart_A_New_Look_at_Maneuvering_Speed.html



And for another take on it, which amplifies things a bit regarding certain control regimes: http://www.flyingmag.com/myth-maneuvering-speed

And for the FAA's definition of maneuvering speed, let's look at FAR 23.335, paragraph c:

(c) Design maneuvering speed VA. For VA, the following applies:
(1) VA may not be less than VS√n where—
(i) VS is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum airplane normal force coefficients, CNA; and
(ii) n is the limit maneuvering load factor used in design.


And 23.1507:


The maximum operating maneuvering speed, VO, must be established as an operating limitation. VO is a selected speed that is not greater than VS√n established in §23.335(c).

 

[dtuuri]

The limiting value for "lift" is not (+4.4g)*(actual gwt). Instead, it's based on the amount of lift the remaining AoA can suddenly generate. When that "tug" on the lift struts is +4.4g*1670, the max limit is reached. Both weights you have listed will generate the maximum allowable lift at the speeds given and yield +4.4g*1670.

dtuuri

Umm... I'd like to retract this. Sorry. In my attempt to bridge across to the OP's way of thinking I missed the fact the "limit" is a load factor, not a lift value. Please disregard.

dtuuri

 

[dmspilot]

Maneuvering speed is all about stalling the airplane before it breaks, folks. At lower gross weights it doesn't stall as readily under heavy G loads, so Va is lower.

I believe this states what everyone already knows and doesn't answer the question.

If for example the wings of a 3,000 lb airplane at 3.8 G are forced to endure 11,400 pounds of load, any more and it would break; why is the 3.8 G the limiting number and not the 11,400 lb.

 

[Dan Thomas]

I believe this states what everyone already knows and doesn't answer the question.

If for example the wings of a 3,000 lb airplane at 3.8 G are forced to endure 11,400 pounds of load, any more and it would break; why is the 3.8 G the limiting number and not the 11,400 lb.

The idea is that if we pull 3.8G at gross at Va, the wing will stall and unload itself. Pull back hard on the elevator, say, and the airplane's path will change slowly enough that the AoA increases to the stall angle at the load limit.

A lighter airplane (less than gross) will change its flight path more quickly and so the AoA doesn't reach stall angle as soon, so the wing can load up beyond the load limit before it stalls, if it stalls at all. Remember that AoA is the relative wind, not some line parallel with the ground, and moving upward reduces the AoA. A lighter airplane therefore has a lower Va, so that the stall speed, which is increasing with the load factor, meets that lower Va so it stalls at the load limit.

Read that Flying article, where the writer describes the all-too-common loss of control in IMC and the airplane spirals out of the cloud deck at high airspeed, sees the ground coming, and pulls back. The load limit is easily exceeded and the wings (or often the tail) come off, because the airspeed is so far above the accelerated stall speed.

 

[Cooter]

The idea is that if we pull 3.8G at gross at Va, the wing will stall and unload itself. Pull back hard on the elevator, say, and the airplane's path will change slowly enough that the AoA increases to the stall angle at the load limit.

A lighter airplane (less than gross) will change its flight path more quickly and so the AoA doesn't reach stall angle as soon, so the wing can load up beyond the load limit before it stalls

That's an interesting way of thinking about it. One thing to remember, the aircraft at a higher gross weight is already flying at a higher AOA (for the same airspeed). It is closer to the critical AOA than the lighter aircraft.

I could be mistaken but I think the "fault" of the OPs understanding was that the published load factor limit was the ACTUAL load limit for both weights, but in reality it is probably only true for the higher weight.

 

[Dan Thomas]

The hypothetical 3000-lb. airplane at 3.8G exerts the 11,400 pounds on the wing, which is designed to take that. At 2700 pounds gross weight the same 11,400 is the limit, but it will take 4.2G to do that. As you say, the 3.8G applies to a gross-weight airplane, which is how the FAA defines strength requirements. They're not going to ask for some load limit strength based on less than gross weight.

And those load ratings are also affected by a 1.5 fudge factor. FAR 23.303:

Unless otherwise provided, a factor of safety of 1.5 must be used.



Now, that doesn't mean that we can load up the airplane safely past gross weight. It also doesn't mean we can go out and pull 5.7G. It means that the FAA wants the engineers to avoid being so obsessed with weight and cost that they design a structure that will fail at 3.8G in the lab. There are inevitably variations in the strengths of materials and quality of workmanship. There is inevitably some loss of strength due to corrosion and fatigue. That 1.5 factor covers those things to a large degree, but we often find corroded items whose strength is pretty much gone. Maintenance is still important to safe flight.

 

[Dan Thomas]

Unfortunately the additional requirements for seats, floors, and engine mounts don't exist in CAR3, which is what many of our light airplanes were certificated under.

The parallel requirement in CAR 3 calls for 6.6G for Normal and Utility categories and 9G for aerobatic. Still stronger than the load factor. For seats, CAR 3, page 22, here: http://specialcollection.dotlibrary.dot.gov/Document?db=DOT-CARS&query=(select+64)

...calls for 3.0 upward, 9.0 forward, and 1.5 sideways load factors, all multiplied by 1.33 (CAR 3.386 and 3.390). That give us 4.0 upward, 12 forward, and 2 sideways, respectively.

Maneuvering speeds won't have much to do with seat or engine mount strengths. It's concerned with not pulling the wings or tail off.

 

[dmspilot]

You still didn't answer the question. If the wing is strong enough to support 3000 pounds at 3.8G, it is strong enough to support 2700 pounds at 4.2G (as you said yourself), or 7.6G at 1500 pounds, and so on. So why do we need to limit it to 3.8G at lighter weights? Why not limit it to 11,400 pounds?

 

[Landshark]

If the wing is rated to handle 3.8g at max gross, and could support more g at lower weights, that would make a whole lot of sense and put an end to my confusion. I still wouldn't go around pulling more g at lower weights. That wing loading is the limiting factor is an assumption that, if wrong, could lead to a very dangerous situation.

Why the max loading figure is a measurement of g at max gross weight and not an ultimate lb load may be the same reason that stall speeds are measured at max gross. It is like a worst case scenario and you'll be fine if you at least follow that limitation to the T.

Which got me to thinking about the effect weight has on certain V speeds. Using the trick that I learned earlier in this thread, I started thinking about how the numbers may change when I am flying solo. Vs0 is 35 KIAS at max gross. If I use my solo weight and plug it in (sqrt(1470/1670) = X/35, the new Vs0 (X) is 32.8 KIAS. I believe this is valid for Vs0, Vs1, Vx, Vy, and of course Va. I cannot do this for Vfe, Vno, or Vne. Here are the new speeds for when I fly solo (1470 lbs).

(sqrt(1470/1670) = X/60 <----you can replace 60 with any approach speed, Vx, Vy etc.

approach speed ................at 1670 lbs = 60 KIAS, at 1470 lbs = 56 KIAS
short field approach speed at 1670 lbs = 54 KIAS, at 1470 lbs = 51 KIAS
Vx speed .........................at 1670 lbs = 55 KIAS, at 1470 lbs = 52 KIAS
Vy speed .........................at 1670 lbs = 67 KIAS, at 1470 lbs = 63 KIAS

What do you guys think?

 

[Dan Thomas]

You still didn't answer the question. If the wing is strong enough to support 3000 pounds at 3.8G, it is strong enough to support 2700 pounds at 4.2G (as you said yourself), or 7.6G at 1500 pounds, and so on. So why do we need to limit it to 3.8G at lighter weights? Why not limit it to 11,400 pounds?

Because the pilot, who often already resists learning any more than he has to in order to get that license, would have to figure out the load limit for every flight, and for every weight change in flight as the fuel burned. 3.8G is a LOT of G loading and is very unlikely to be encountered if the airplane is flown sensibly. A 60 degree banked turn generates 2G, a 70 degree banked turn is 3G, and those are pretty steep. Most pilots won't go much beyond 45 degrees. At 70 degrees most lightplanes don't have the power to maintain level flight anyway. The drag is too high.

Most pilots who pull hard enough to stall the airplane and spin and crash do it with some stupid thing like a buzz job with a sharp pull-up that raises the stall speed until it meets the airspeed and it's all over and he doesn't know what he did. I cringe every time I see some guy do that, especially if he throws a turn into the maneuver at the same time. Looks cool, but it can kill quickly.

Too many pilots don't understand the relationship between load factor and angle of attack and stall speed. It's tough enough to get them to learn that, without asking them to calculate the permissible load factor for a particular flight. Va for the current weight is enough. If I thought my students were out there deliberately pulling 3.8G for the fun of it I'd be ticked. Sooner or later one of them will inadvertently exceed that and maybe weaken the airplane so it kills someone later on. Or wrinkle it and write it off. Besides, as the Flying article says, there are other ways to break the airplane within Va.

 

[MAKG1]

If the wing is rated to handle 3.8g at max gross, and could support more g at lower weights, that would make a whole lot of sense and put an end to my confusion. I still wouldn't go around pulling more g at lower weights. That wing loading is the limiting factor is an assumption that, if wrong, could lead to a very dangerous situation.

Why the max loading figure is a measurement of g at max gross weight and not an ultimate lb load may be the same reason that stall speeds are measured at max gross. It is like a worst case scenario and you'll be fine if you at least follow that limitation to the T.

Which got me to thinking about the effect weight has on certain V speeds. Using the trick that I learned earlier in this thread, I started thinking about how the numbers may change when I am flying solo. Vs0 is 35 KIAS at max gross. If I use my solo weight and plug it in (sqrt(1470/1670) = X/35, the new Vs0 (X) is 32.8 KIAS. I believe this is valid for Vs0, Vs1, Vx, Vy, and of course Va. I cannot do this for Vfe, Vno, or Vne. Here are the new speeds for when I fly solo (1470 lbs).

(sqrt(1470/1670) = X/60 <----you can replace 60 with any approach speed, Vx, Vy etc.

approach speed ................at 1670 lbs = 60 KIAS, at 1470 lbs = 56 KIAS
short field approach speed at 1670 lbs = 54 KIAS, at 1470 lbs = 51 KIAS
Vx speed .........................at 1670 lbs = 55 KIAS, at 1470 lbs = 52 KIAS
Vy speed .........................at 1670 lbs = 67 KIAS, at 1470 lbs = 63 KIAS

What do you guys think?

Careful....

That square root thing operates on calibrated airspeed.

While there isn't a whole lot of difference at moderate speeds, there is at low speeds. 51 KIAS approach is too slow.

 

[dmspilot]

Because the pilot, who often already resists learning any more than he has to in order to get that license, would have to figure out the load limit for every flight, and for every weight change in flight as the fuel burned...

Too many pilots don't understand the relationship between load factor and angle of attack and stall speed. It's tough enough to get them to learn that, without asking them to calculate the permissible load factor for a particular flight. Va for the current weight is enough.

But you wouldn't need to calculate anything. A given airplane will stall under a load of 11,400 pounds at the same airspeed regardless of its actual weight, so if Va was based on a load limit instead of a load factor limit no calculation is necessary. It would actually be easier on the pilot.

The rest of your post about low-altitude buzz jobs and pulling 3.8G for fun is totally irrelevant to the discussion.

 

[Dan Thomas]

The rest of your post about low-altitude buzz jobs and pulling 3.8G for fun is totally irrelevant to the discussion.

The Va limit is all about accelerated stalls. And accelerated stalls, I have found, are not clearly understood by too many pilots and so there are too many silly accidents.

Gust loads cause G loading which causes higher AoA which causes accelerated stalls. That's one reason why we have Va.

 

[Landshark]

Careful....

That square root thing operates on calibrated airspeed.

While there isn't a whole lot of difference at moderate speeds, there is at low speeds. 51 KIAS approach is too slow.

On my particular airplane, the calibration error is much less at slow speeds. Only at 140 knots with flaps up is there a 4 knot error. Here is a part of the airspeed calibration table.


*Flaps at 30 degrees

KIAS.....50......60
KCAS....51......61


With what you said about different levels of error at different speeds and the fact that there were specific airspeed calibration tables for different flap configurations in the POH, something occurred to me in terms of calibration error.

The main difference in all those situations was the different AOA and corresponding angle between the relative wind and the pitot tube. The amount of calibration error must be influenced by the angle between the relative wind and the pitot tube. That would also explain why the airspeed indicator is not accurate in a forward slip (besides the interference from the airframe in a lefthand forward slip).

 

[dtuuri]

The amount of calibration error must be influenced by the angle between the relative wind and the pitot tube. That would also explain why the airspeed indicator is not accurate in a forward slip (besides the interference from the airframe in a lefthand forward slip).

I think the pressure change at the static source is the bigger culprit.

As for what I think about adjusting approach speed as you did in your table, I think it's ok because the standard "1.3 x power off stalling speed" is included in the adjusted value as well.

As for Vx and Vy, technically, I'm not so sure although it's close enough for government work. Vy, for instance, occurs where there is the most excess horsepower--and I think that increases with airspeed in a unique way. Same goes for Vx, except in that case it's excess thrust. These are interesting questions, though.

dtuuri

 

[Landshark]

As for Vx and Vy, technically, I'm not so sure although it's close enough for government work. Vy, for instance, occurs where there is the most excess horsepower--and I think that increases with airspeed in a unique way. Same goes for Vx, except in that case it's excess thrust. These are interesting questions, though.

dtuuri

I searched some of the keywords in your post and found a pretty informative article by Rod Machado about Vx, Vy and different altitudes. Too bad it wasn't about different weights

It explained that Vy happens at the airspeed where there is the largest gap between the power available curve and the power required curve. The power available curve shouldn't change with different weights. For a lower airplane weight, if the power required curve maintains the same shape, but is just lowered at every point, then Vy should actually remain at the same speed as before. You would just have more excess power and a greater climb rate at that speed.

If the power required curve changes shape with less weight, I guess you would have to look at both the new power required and power available curves to determine Vy.

 

[dtuuri]

I searched some of the keywords in your post and found a pretty informative article by Rod Machado about Vx, Vy and different altitudes. Too bad it wasn't about different weights

You may be interested in Fig 2.2 on page 100 of Aerodynamics for Naval Aviators.

dtuuri

 

[Landshark]

Thank you for that info. For the lower weight, it looks like the curve flattens out near the left side. The airspeed for minimum power required is also shifted to the left. At first glance, that would make you wanna think that Vy is lowered the same amount. But in my case, available power also decreases as the airspeed goes down.

An educated guess would be that Vy does go down with decreasing weight, just not as much as you might expect. I suppose I could graph a power required curve given a couple numbers like Cd, Cl, and some square footage areas, but the available power curve (and propeller efficiency) is beyond me.

 

[kzadora]

A simple answer to your original question, Calvin, is that maneuvering speed is representing a limit on acceleration.

A wing travelling at the same speed and angle of attack produces the same amount of lift. This does not change with the airplane total weight but does change with the airspeed and angle of attack. As the airplane weight decreases, the maximum acceleration you can get out of a wing before it stalls increases. So to keep the maximum acceleration in check, the maximum speed (maneuvering speed) must decrease with weight.

 

[noahfong]

Some of the folks alluded to it, but to address your curiosity as to why a lower weight results in a lower Va...

Some facts:

1. A wing stalls at a critical angle of attack. Weight and airspeed has nothing to do with it.

2. Lift is a function of angle of attack and airspeed.

So, an airplane maintaining altitude at a given airspeed will have a larger angle of attack when it is heavier. This is because, for the wing to generate more lift without increasing speed, it needs an increased angle of attack.

So, a heavier airplane will be at a higher angle of attack than a lighter one at the same airspeed. let's take it further....

Va is calculated based on an angle of attack such that the wing will stall (reach the critical AOA) before the load is high enough to break something. Let's say this is AOA is achieved at 90kts for an airplane weighing 2000lbs. If we lighten the airplane to 1600lbs, this same angle of attack may be achieveable only if the airplane is SLOWED to 85kts. Remember, less lift is needed and if we want to hold the AOA constant, we will have to reduce the airspeed.

That's why Va goes DOWN with weight.

Remember - weight and airspeed has nothing to do with the wing stalling. It's all AOA. Your calculations, although interesting, is really besides the issue of Va.

 

[Landshark]

It is interesting that almost everyone thinks about it differently than myself, even though we end up at the same result. I think of Va as the slowest speed to achieve the limit load factor(g). Load factor is a function of lift and weight. So for Va, I think of how much lift the wing is capable of producing at Va, then divide that by the weight to arrive at the g load factor limit.

Clmax(critical AOA) at Va = max lift at that speed. Max lift/weight = load factor limit. Pull back harder on the stick and you would be met with a stall and less lift, therefore it is safe to do so at Va. Any higher velocity would be capable of more lift, so you have to be careful with inputs above Va. Higher weight needs more lift(and speed) to get the same g force.

In both ways of looking at it, we consider critical AOA. Since almost everyone looks at it differently than myself, I wonder if there is an advantage to thinking about it in terms of AOA instead of load factor. Maybe some conceptual advantage that makes understanding related subjects easier or more intuitive.

 

[Cooter]

It is interesting that almost everyone thinks about it differently than myself, even though we end up at the same result. I think of Va as the slowest speed to achieve the limit load factor(g). Load factor is a function of lift and weight. So for Va, I think of how much lift the wing is capable of producing at Va, then divide that by the weight to arrive at the g load factor limit.

Va would coincide with Clmax(critical AOA) X velocity = max lift at that speed. Pull back harder on the stick and you would be met with a stall and less lift, therefore it is safe to do so at Va. Any higher velocity would be capable of more lift, so you have to be careful with inputs above Va. Higher weight needs more lift(and speed) to get the same g force.

In both ways of looking at it, we consider critical AOA. Since almost everyone looks at it differently than myself, I wonder if there is an advantage to thinking about it in terms of AOA instead of load factor. Maybe some conceptual advantage that makes understanding related subjects easier or more intuitive.

I think of it as G available. For any given airspeed(assuming weight is constant), you only have so much G available before Critical AOA is reached. The lighter you are, the more G is available for any given speed.

 

[MAKG1]

I think of it as G available. For any given airspeed(assuming weight is constant), you only have so much G available before Critical AOA is reached. The lighter you are, the more G is available for any given speed.

This doesn't work.

Straight and level, the "G available" is always max - 1. But, a heavier airplane at the same speed (like Noah said) has a larger angle of attack.

The upshot is that AoA is not acceleration, it's not G's, it's not weight, it's not pitch angle (common error there), and it's not speed. All of them can change the angle of attack, but it's not correct to think of them as a substitute for AoA.

 

[Cooter]

This doesn't work.

Straight and level, the "G available" is always max - 1. But, a heavier airplane at the same speed (like Noah said) has a larger angle of attack.

Sure it does. For any given airspeed and weight, there is a certain amount of G available for turning(or performance). Why do you think fighter pilots used to punch off their drop tanks before an engagement? I'm not sure what you mean by G available is always 1, unless we're talking about different things.

In the post I was responding to, "it" was referring to Va, not AoA. Va references maximum load factor, or G, and airspeed.

 

[jaymark6655]

I read all the other post and it made my head hurt.

The way I look at it, Va is the speed at which you can move the control one direction through its full deflection and not bend something. The reason is because of the resulting Gs, which is just acceleration. If you are lighter you can "change direction" (acceleration or G) faster (higher resulting G), than if you are heavier.

 

[Dan Thomas]

I read all the other post and it made my head hurt.

The way I look at it, Va is the speed at which you can move the control one direction through its full deflection and not bend something. The reason is because of the resulting Gs, which is just acceleration. If you are lighter you can "change direction" (acceleration or G) faster (higher resulting G), than if you are heavier.

Google '"accelerated stalls" and read everything you can find on the subject. This is one area that too few pilots understand, and it's killing them.

 

[jaymark6655]

Google '"accelerated stalls" and read everything you can find on the subject. This is one area that too few pilots understand, and it's killing them.

Thanks. I know what they are and have experienced some. Just think most of this seemed to veer way of course.

Va decrease with weight because a smaller AOA is needed to fly than if you were heavier. Remember an airfoil will stall at a certain AOA every time. Anytime you double AOA you almost double Gs. At light weight if your AOA is 3 degrees for level flight at X speed, anything that changes it to 6 just doubled your Gs and 12 will be close to 4. At gross if your AOA for level flight at X speed is 5 you would have to increase to 10 to get 2Gs and 20 for 4 Gs, but by then the airfoil has stalled and is no longer loaded. Lighter weight also means less inertia so the movement from 3 degrees to 6 can happen a lot faster than it would if you were at gross. If you slow from X speed to Y when the aircraft is light, the AOA has to increase to maintain level flight and the AOA will move past the Critical AOA before reaching the G limit, removing the load.

Accelerated Stall = something other than level flight that increase the load demand of the wings causing the need for an increased AOA. This increase in AOA to keep up with the demand for lift due to turning for example finally exceeds critical AOA causing a stall even though airspeed might be 80 and stall speed for the aircraft is 48 knots.

 

[dmspilot]

OP asked why we care about acceleration (Gs) and not load (e.g. pounds) but everyone wants to answer a different question.

The easiest explanation why Va decreases with weight is because it is a function of stall speed, which decreases with weight. It is (usually) the stall speed at the limit load factor.

e.g. Va = Vs * √3.8

 

[Dan Thomas]

x

The easiest explanation why Va decreases with weight is because it is a function of stall speed, which decreases with weight. It is (usually) the stall speed at the limit load factor.

e.g. Va = Vs * √3.8

Of course. So why didn't you say that in your previous posts? Everybody was going on about pounds of load on the wings and load limits and all sort of stuff, everything except the reall issue: the stalling speed at load factor limit and different weights and at what airspeed that happens.

 

[Cooter]

OP asked why we care about acceleration (Gs) and not load (e.g. pounds) but everyone wants to answer a different question.

The OP made a faulty assumption about load factors at less than gross weight. His question asked where the fault in his reasoning was and it was pointed out to him, so his question was answered. The rest was general discussion about load factors and AoA where everybody had a different way of looking at it.