Load factor limitations
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Load factor limitations
Hi all,
I've been helping my girlfriend study for her CPL(A) Performance exam the last few days and I've gone and confused myself.
Can someone please help me clarify a few things:
(a) Is it the limit load factor only, or a combination of the limit load factor and weight that determines the structural limits of an aircraft design.
My logic is this. Load factor is just another word for acceleration. If you put the aircraft under a load factor of 2 you are essentially accelerating the aircraft with a force of 2x gravity. If you have a 1000kg aircraft in a 60 degree level turn you require a load factor of 2.0g the wings must therefore provide a force of 2000kg. However if you overload the same aircraft by 500kg to 1500kg then the same wings have to provide a force of 3000kg to maintain the turn. Surely the heavier aircraft is put under greater structural stress. Is it the case then that two aircraft of different weight can be put under the same load factor but have different degrees of structural stress imparted on the structure. Is this the reason for the utility category limits of some aircraft only being valid below certain weights.
(b) Why is it that Va increases with increased weight.
The only thing that I can think of is that if the aircraft is heavier the lift required from the wing to generate the same load factor (lift/weight) increases thus a higher speed and greater lift for given AoA will have less effect on the load factor. Which again leads to the first question regarding structural limits.
Hope I'm not way off here!
Cheers
J
I've been helping my girlfriend study for her CPL(A) Performance exam the last few days and I've gone and confused myself.
Can someone please help me clarify a few things:
(a) Is it the limit load factor only, or a combination of the limit load factor and weight that determines the structural limits of an aircraft design.
My logic is this. Load factor is just another word for acceleration. If you put the aircraft under a load factor of 2 you are essentially accelerating the aircraft with a force of 2x gravity. If you have a 1000kg aircraft in a 60 degree level turn you require a load factor of 2.0g the wings must therefore provide a force of 2000kg. However if you overload the same aircraft by 500kg to 1500kg then the same wings have to provide a force of 3000kg to maintain the turn. Surely the heavier aircraft is put under greater structural stress. Is it the case then that two aircraft of different weight can be put under the same load factor but have different degrees of structural stress imparted on the structure. Is this the reason for the utility category limits of some aircraft only being valid below certain weights.
(b) Why is it that Va increases with increased weight.
The only thing that I can think of is that if the aircraft is heavier the lift required from the wing to generate the same load factor (lift/weight) increases thus a higher speed and greater lift for given AoA will have less effect on the load factor. Which again leads to the first question regarding structural limits.
Hope I'm not way off here!
Cheers
J
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Upward vertical gusts will increase the angle of attack of a wing in flight. Downward vertical gusts will decrease the angle of attack in a similar manner. When flying through turbulence, these random changes in angle of attack produce random changes in the lift force.
Load factor is equal to lift divided by weight. In straight and level flight the lift is equal to the weight and the load factor is one. Any sudden increase in lift caused by an upward gust will increase the angle of attack. This will increase both the lift and the load factor. This increased load factor is called the gust load, or gust response.
Factors that determine the magnitude of the gust load include the vertical speed of the gust and the degree to which the angle of attack is increased by it. If for example the angle of attack is doubled then the load factor will also be doubled from one to two, so the gust load will be two. At any given airspeed the number of degrees by which the angle of attack is increased is determined by the vertical speed of the gust. But the fraction by which the angle of attack increases also depends upon the angle of attack prior to the gust.
If for example, if the initial angle of attack is one degree, and a gust increases it by one degree, then the angle of attack and the load factor have doubled. But if the initial angle of attack is two degrees and the same gust increases it by the same one degree then the increase is only 50% so the gust load is 1.5. So any factor that decreases the initial angle of attack will increase the gust load for any given gust strength. The factors that decrease the initial angle of attack include decreasing aircraft weight and increasing aircraft speed. So gust load increases when the weight decreases.
Load factor is equal to lift divided by weight, so these sudden changes in lift will cause changes in the load factor. Assuming that the initial condition is straight and level flight with a load factor of 1, then for JAR examination purposes the new load factor in any gust can be calculated using the equation:
Load factor in a gust = (New lift / (Old lift)
So load factor in a gust = (New CL) / (Old CL)
Where
New CL is CL in the gust
Old CL is the CL prior to the gust
VA is the airspeed at which the aircraft will stall at the limiting load factor. So in a sense, VA is just another stalling speed. Stalling speeds are proportional to the square root of weight, so all stalling speeds increase with increasing weight.
Students are often puzzled by the fact that VA decreases when aircraft weight decreases. They often ask “well we can stall it at XXX knots without overloading the aircraft at a mass of 75000 kg, so why can we not do so at a mass of only 50000 kg?”
The answer is that at lower weights, the same airspeed will enable us to achieve greater load factors before we stall. Although the wings might not break off, some of the other components may start to do so. The weight of the aircraft may have decreased, but the weights of individual items like engines will not have changed. So imposing a greater g will overload their mountings.
Load factor is equal to lift divided by weight. In straight and level flight the lift is equal to the weight and the load factor is one. Any sudden increase in lift caused by an upward gust will increase the angle of attack. This will increase both the lift and the load factor. This increased load factor is called the gust load, or gust response.
Factors that determine the magnitude of the gust load include the vertical speed of the gust and the degree to which the angle of attack is increased by it. If for example the angle of attack is doubled then the load factor will also be doubled from one to two, so the gust load will be two. At any given airspeed the number of degrees by which the angle of attack is increased is determined by the vertical speed of the gust. But the fraction by which the angle of attack increases also depends upon the angle of attack prior to the gust.
If for example, if the initial angle of attack is one degree, and a gust increases it by one degree, then the angle of attack and the load factor have doubled. But if the initial angle of attack is two degrees and the same gust increases it by the same one degree then the increase is only 50% so the gust load is 1.5. So any factor that decreases the initial angle of attack will increase the gust load for any given gust strength. The factors that decrease the initial angle of attack include decreasing aircraft weight and increasing aircraft speed. So gust load increases when the weight decreases.
Load factor is equal to lift divided by weight, so these sudden changes in lift will cause changes in the load factor. Assuming that the initial condition is straight and level flight with a load factor of 1, then for JAR examination purposes the new load factor in any gust can be calculated using the equation:
Load factor in a gust = (New lift / (Old lift)
So load factor in a gust = (New CL) / (Old CL)
Where
New CL is CL in the gust
Old CL is the CL prior to the gust
VA is the airspeed at which the aircraft will stall at the limiting load factor. So in a sense, VA is just another stalling speed. Stalling speeds are proportional to the square root of weight, so all stalling speeds increase with increasing weight.
Students are often puzzled by the fact that VA decreases when aircraft weight decreases. They often ask “well we can stall it at XXX knots without overloading the aircraft at a mass of 75000 kg, so why can we not do so at a mass of only 50000 kg?”
The answer is that at lower weights, the same airspeed will enable us to achieve greater load factors before we stall. Although the wings might not break off, some of the other components may start to do so. The weight of the aircraft may have decreased, but the weights of individual items like engines will not have changed. So imposing a greater g will overload their mountings.
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A very nice post Keith, well worthy of a 'cut and paste' for all aspiring pilots.
I hope that my following comments do not 'pollute' Keith's fine words, they address one point raised in the original post.
If you are flying a Normal category aircraft, it is subject to a 2.5G limit for all normal operations. Structural deformation will occur at 2.5G, but the airframe will return to normal once returned to 1G flight. There is a 150% margin bulit into this 2.5G limit, that is, the aircraft may be subject to 3.75G before PERMANENT structural deformation occurs. Beyond this, structural failure is a distinct possibility. The additional 50% margin is NEVER to be used deliberately, it is there for your protection as a buffer if further inadvertant load occur when operating at 2.5G.
When a Normal category aircraft is certified, it is capable of meeting a 2.5G/3.75G requirement AT THE MAXIMUM PERMITTED OPERATING WEIGHT. Typically, no higher 'G' limit is published for lower weights (Most limitations are for the worst case). For example, an aircraft limited to 1000 Kg Maximum operating Weight, has sufficient structural integrity to withstand operations to 2500 Kg / 3750 Kg. At lower weights the aircraft can withstand higher 'G' forces before reaching the limit loads, but this is NOT utilised for normal operations. Some Normal category aircraft are permitted limited aerobatics at lower limiting weights for this reason.
I have lingering bitter memories of pilot abuse of aircraft load limits. Many years ago, a close friend died following structural failure in a Normal category aircraft. He only encountered moderately turbulent conditions which lead to the structural failure. It transpired that another pilot had previously been regularly been doing aerobatics in the aircraft which (unfortunately) did not kill him, but applied cumulative stresses to the aircraft which finally failed many pilots later.
If you overstress an aircraft - REPORT IT, and save subsequent user's lives.
Regards,
Old Smokey
I hope that my following comments do not 'pollute' Keith's fine words, they address one point raised in the original post.
If you are flying a Normal category aircraft, it is subject to a 2.5G limit for all normal operations. Structural deformation will occur at 2.5G, but the airframe will return to normal once returned to 1G flight. There is a 150% margin bulit into this 2.5G limit, that is, the aircraft may be subject to 3.75G before PERMANENT structural deformation occurs. Beyond this, structural failure is a distinct possibility. The additional 50% margin is NEVER to be used deliberately, it is there for your protection as a buffer if further inadvertant load occur when operating at 2.5G.
When a Normal category aircraft is certified, it is capable of meeting a 2.5G/3.75G requirement AT THE MAXIMUM PERMITTED OPERATING WEIGHT. Typically, no higher 'G' limit is published for lower weights (Most limitations are for the worst case). For example, an aircraft limited to 1000 Kg Maximum operating Weight, has sufficient structural integrity to withstand operations to 2500 Kg / 3750 Kg. At lower weights the aircraft can withstand higher 'G' forces before reaching the limit loads, but this is NOT utilised for normal operations. Some Normal category aircraft are permitted limited aerobatics at lower limiting weights for this reason.
I have lingering bitter memories of pilot abuse of aircraft load limits. Many years ago, a close friend died following structural failure in a Normal category aircraft. He only encountered moderately turbulent conditions which lead to the structural failure. It transpired that another pilot had previously been regularly been doing aerobatics in the aircraft which (unfortunately) did not kill him, but applied cumulative stresses to the aircraft which finally failed many pilots later.
If you overstress an aircraft - REPORT IT, and save subsequent user's lives.
Regards,
Old Smokey
Indeed,...and remember that VTP/RA is NOT fully protective,... the wings staying on is...slow way down for the extreme stuff
Old Smokey check your PM about the 'hydraulic non return valve' [now I remember the word]
Old Smokey check your PM about the 'hydraulic non return valve' [now I remember the word]
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PMs checked PA, E-mail to follow. I think that I see where you're coming from, you may need to refer to "English for Dummies" (OS = Dummy) before I can fully comprehend!
Thanks for the feedback.
Regards,
Old Smokey
Thanks for the feedback.
Regards,
Old Smokey
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Structural Load Limit
I have two questions:
- Why is it that an extension of flaps reduces the structural load limit?
- Why is the negative load limit zero (0) once flaps are extended?
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If an airplane has a 2.5G load limit (ignoring for now any safety factors), the WING will not likely failwith any load up to 2.5 x max gross takeoff weight. HOWEVER, individual components may only be stressed to 2.5G, so a 3G stress will break them.
An example I read in a magazine many years ago was the battery shelf on a light SE airplane. If the airplane was limited to 4G, then the battery shelf was likely designed to a 4G limit. Even if the wings could withstand 5G at light gross weights, that battery shelf would break, and may cause catastrophic consequential damage.
An example I read in a magazine many years ago was the battery shelf on a light SE airplane. If the airplane was limited to 4G, then the battery shelf was likely designed to a 4G limit. Even if the wings could withstand 5G at light gross weights, that battery shelf would break, and may cause catastrophic consequential damage.
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An example I read in a magazine many years ago was the battery shelf on a light SE airplane. If the airplane was limited to 4G, then the battery shelf was likely designed to a 4G limit. Even if the wings could withstand 5G at light gross weights, that battery shelf would break, and may cause catastrophic consequential damage.
BTW, load limit is a term used specifically for aircraft STRUCTURE.
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First, reread the part where said, "(ignoring for now any safety factors)"...
Then, in the case of a transport category aircraft, the "anticipated load limit of the aircraft" is 2.5G. Therefore, any components are designed to withstand 2.5G; that means a component support structure is designed to carry 2.5x the normal weight of the components it supports (e.g., a battery rack, as noted before).
All that is notwithstanding the fact that the WING is designed to carry 2.5x the max gross takeoff weight of the aircraft. That means that even though the aircraft is not certified to more than 2.5G, at lower gross weights the wing can withstand a higher load factor without increased chance of WING failure. However, since all components within the airframe are only required to be designed to the same 2.5G load factor, those components are at higher risk of failure if more than 2.5G are applied.
This concept is illustrated clearly by many military fighter aircraft, where max allowed G load varies significantly with gross weight, usually targeting a constant gross weight x load factor figure. At max T/O weight, the max allowed load factor may be 3 or 4.5G, where the max allowed load factor at lighter weight may be 6.5 or 9G. In that type of aircraft, the components are designed to the max allowed load factor at any time (e.g., 9G), not the max load factor at max gross weight (e.g., 4.5G).
The bottom line is that one can NOT assume that it is safe to fly an airplane at a load factor greater than its certified load factor.
As for the differences when the flaps are extended, again the flap carriage structure is only required to be designed to the lesser (0 to +2.0G) limits. With the flaps retracted, the flap carriage structures are not stressed the same way as when the flaps are extended.
Then, in the case of a transport category aircraft, the "anticipated load limit of the aircraft" is 2.5G. Therefore, any components are designed to withstand 2.5G; that means a component support structure is designed to carry 2.5x the normal weight of the components it supports (e.g., a battery rack, as noted before).
All that is notwithstanding the fact that the WING is designed to carry 2.5x the max gross takeoff weight of the aircraft. That means that even though the aircraft is not certified to more than 2.5G, at lower gross weights the wing can withstand a higher load factor without increased chance of WING failure. However, since all components within the airframe are only required to be designed to the same 2.5G load factor, those components are at higher risk of failure if more than 2.5G are applied.
This concept is illustrated clearly by many military fighter aircraft, where max allowed G load varies significantly with gross weight, usually targeting a constant gross weight x load factor figure. At max T/O weight, the max allowed load factor may be 3 or 4.5G, where the max allowed load factor at lighter weight may be 6.5 or 9G. In that type of aircraft, the components are designed to the max allowed load factor at any time (e.g., 9G), not the max load factor at max gross weight (e.g., 4.5G).
The bottom line is that one can NOT assume that it is safe to fly an airplane at a load factor greater than its certified load factor.
As for the differences when the flaps are extended, again the flap carriage structure is only required to be designed to the lesser (0 to +2.0G) limits. With the flaps retracted, the flap carriage structures are not stressed the same way as when the flaps are extended.
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Thanks for all the responses,
The confusion is gone now, and as my 2 cents, limits in the flight manual are LIMITS, I'm sure a lot of boffins have put a great many hours into coming up with them so I won't be in any hurry to exceed them. Unless of course the houses are getting bigger much too quickly.
J
The confusion is gone now, and as my 2 cents, limits in the flight manual are LIMITS, I'm sure a lot of boffins have put a great many hours into coming up with them so I won't be in any hurry to exceed them. Unless of course the houses are getting bigger much too quickly.
J
