1.5 Factor Of Safety - FAR 25.303
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1.5 Factor Of Safety - FAR 25.303
Hi all,
Does the Factor of Safety of 1.5 under FAR Part 25.303 apply to speeds or just G-loads?
eCFR ? Code of Federal Regulations
In other words, is there a factor of safety of 1.5 above the speed Vne or Vd as well?
In further terms, looking at the following VG diagram, there is an "Overspeed margin" between Vne and the end of the flight envelope.
Does the "Overspeed margin" represent a Factor of 1.5 as well as the G-loadsl?
Thanks in advance.
Does the Factor of Safety of 1.5 under FAR Part 25.303 apply to speeds or just G-loads?
eCFR ? Code of Federal Regulations
In other words, is there a factor of safety of 1.5 above the speed Vne or Vd as well?
In further terms, looking at the following VG diagram, there is an "Overspeed margin" between Vne and the end of the flight envelope.
Does the "Overspeed margin" represent a Factor of 1.5 as well as the G-loadsl?
Thanks in advance.
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Thank you for your reply bigduke...
So, if I were to state -
"a 1.5 factor of safety (yellow) is included for ALL loads that planes must withstand, not just the g-loads shown on a Vg diagram" -- with regard to the VG diagram posted above, would that be true or false?
So, if I were to state -
"a 1.5 factor of safety (yellow) is included for ALL loads that planes must withstand, not just the g-loads shown on a Vg diagram" -- with regard to the VG diagram posted above, would that be true or false?
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FAR 25 defines a set of loads (not just g loads) that (nominally) may be expected once in the life of an airplane. These are limit loads. The structure must be demonstrated by test to be able to withstand 1.5 times these loads.
These limit loads are defined against structural design speed requirements Va,Vb,Vc and Vd - there are no such speeds as Vno and Vne in FAR 25.
Va,b,c,d speeds are defined in 25.335
The requirements associated with Vd/Md have to be demonstrated by flight test if Md is greater than 0.8M. FAR 25.629 defines an enlarged flight envelope beyond Vd/Md within which certain aeroelastic criteria must be met.
Obviously any margins required over the flight demonstrated limits have to be "proved" by calculation. The aircraft must shown to by calculation be free of flutter, control reversal etc. up to 1.15 Vd/Md (unless this speed would exceed Mach 1.0)
I see this 1.15 factor (was 1.2) as similar in concept to the 1.5 load factor of safety but of course it is a 'paper' margin where the 1.5 factor has to be physically demonstrated.
These limit loads are defined against structural design speed requirements Va,Vb,Vc and Vd - there are no such speeds as Vno and Vne in FAR 25.
Va,b,c,d speeds are defined in 25.335
The requirements associated with Vd/Md have to be demonstrated by flight test if Md is greater than 0.8M. FAR 25.629 defines an enlarged flight envelope beyond Vd/Md within which certain aeroelastic criteria must be met.
Obviously any margins required over the flight demonstrated limits have to be "proved" by calculation. The aircraft must shown to by calculation be free of flutter, control reversal etc. up to 1.15 Vd/Md (unless this speed would exceed Mach 1.0)
I see this 1.15 factor (was 1.2) as similar in concept to the 1.5 load factor of safety but of course it is a 'paper' margin where the 1.5 factor has to be physically demonstrated.
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Hi Owain,
Thank you for your reply. Yes, I agree FAR 25 defines many different types of loads... but would speed be considered a "load" regulated by FAR 25.303?
I always understood that 1.5 factor of safety does not apply to the speeds, as stated by bigduke above.
Yes, I understand the Flutter Free margin (1.15 or 1.2VD) is only based on theoretical calculations and never demonstrated, and it is for equivalent airspeed at both constant mach and altitude (meaning no changes in static or dynamic pressure which might induce flutter at a lower speed).
But the VG envelope is not increased by such a margin due to the fact the VG includes G loads, and therefore would induce flutter above Vd. Correct?
Would I also be correct to say that the "Overspeed Margin" in the above VG diagram is the 10% margin between Vne and Vd as defined under FAR 23.1505?
Or is there some other margin above Vne in a VG diagram under FAR Part 23?
Thank you for your reply. Yes, I agree FAR 25 defines many different types of loads... but would speed be considered a "load" regulated by FAR 25.303?
I always understood that 1.5 factor of safety does not apply to the speeds, as stated by bigduke above.
Yes, I understand the Flutter Free margin (1.15 or 1.2VD) is only based on theoretical calculations and never demonstrated, and it is for equivalent airspeed at both constant mach and altitude (meaning no changes in static or dynamic pressure which might induce flutter at a lower speed).
But the VG envelope is not increased by such a margin due to the fact the VG includes G loads, and therefore would induce flutter above Vd. Correct?
Would I also be correct to say that the "Overspeed Margin" in the above VG diagram is the 10% margin between Vne and Vd as defined under FAR 23.1505?
Or is there some other margin above Vne in a VG diagram under FAR Part 23?
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I think your mistake is assuming that all the values are calculated.
in practise they are not.
what far 23 is all about is setting strength standards that allow for foreseeable gust loads in typical operating environments.
this saves a designer from having to do laborious calculations of gust likelihood and effect.
design to the far strengths for each category and the loads will work out ok in practise.
Vne is 90% of whatever was demonstrated during test flying as Vd. Vd is actually reached in flight during the test. if flutter is not experienced at Vd then that sets the value of Vne.
in your diagram the outer edge of the white and green represents the "limit" load.
the outer edge of the yellow is the "ultimate" load.
as a designer you can expect all loads caused by the pilot to be within the limit load.
between the limit strength and the ultimate strength the pilot can expect damage to occur but not damage that would prove fatal.
the aircraft is typically designed to be just stronger than the calculated "ultimate" strength.
in practise they are not.
what far 23 is all about is setting strength standards that allow for foreseeable gust loads in typical operating environments.
this saves a designer from having to do laborious calculations of gust likelihood and effect.
design to the far strengths for each category and the loads will work out ok in practise.
Vne is 90% of whatever was demonstrated during test flying as Vd. Vd is actually reached in flight during the test. if flutter is not experienced at Vd then that sets the value of Vne.
in your diagram the outer edge of the white and green represents the "limit" load.
the outer edge of the yellow is the "ultimate" load.
as a designer you can expect all loads caused by the pilot to be within the limit load.
between the limit strength and the ultimate strength the pilot can expect damage to occur but not damage that would prove fatal.
the aircraft is typically designed to be just stronger than the calculated "ultimate" strength.
Last edited by dubbleyew eight; 28th Jun 2014 at 14:30.
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Yes, I understand. And in fact these are not my mistakes, rather they are the mistakes made by someone with which I have been having a discussion.
Basically, they are asserting that FAR 25.301, 303, and 305 regulate speeds as loads. And that the 1.5 safety factor defined under 25.303 pertains to speeds.
I have kept trying to tell them that the loads defined in FAR 25.301, 303 and 305 are not speeds, and are good up to Vd, which is defined in FAR 25.305.
For example, there isn't any margin of safety covering a Transport Category aircraft which is accelerating through Vd+50 while pulling 2-3.8 G's (limit load + 1.5 factor of safety).
The person I am having this discussion with seems to think such a maneuver is covered by FAR 25.301, 303, and 305.
Basically, they are asserting that FAR 25.301, 303, and 305 regulate speeds as loads. And that the 1.5 safety factor defined under 25.303 pertains to speeds.
I have kept trying to tell them that the loads defined in FAR 25.301, 303 and 305 are not speeds, and are good up to Vd, which is defined in FAR 25.305.
For example, there isn't any margin of safety covering a Transport Category aircraft which is accelerating through Vd+50 while pulling 2-3.8 G's (limit load + 1.5 factor of safety).
The person I am having this discussion with seems to think such a maneuver is covered by FAR 25.301, 303, and 305.
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Speed as such is not a 'load'. 2.5g at 250kts and a given weight/CG is the same load on the structure as 2.5g at 150 kts and that weight/CG (those speeds chosen to eliminate compressibility effects). The loadings are defined at specific speeds so speed cannot really be also subject to a 1.5 factor.
Not sure what you have in mind here. Equivalent airspeed defines dynamic pressure so Vd at any altitude gives the maximum dynamic pressure there. Flutter gets more likely as dynamic pressure increases so lower speeds are less critical.
I've never heard of anyone checking flutter at speeds above Vd/Md combined with pulling 'g'. The aircraft is of course cleared for flutter at Vd/Md and 1.5g (at least) as that is part of the Vdf/Mdf demonstration. The additional calculated margins are intended to make sure that there is no 'cliff edge' just beyond Vd/Md
Sorry, I thought from your OP that you were asking about Part 25 aircraft. I'm not sufficiently familiar with Part 23 to give you a reliable answer
Yes, I understand the Flutter Free margin (1.15 or 1.2VD) is only based on theoretical calculations and never demonstrated, and it is for equivalent airspeed at both constant mach and altitude (meaning no changes in static or dynamic pressure which might induce flutter at a lower speed).
But the VG envelope is not increased by such a margin due to the fact the VG includes G loads, and therefore would induce flutter above Vd. Correct?
Would I also be correct to say that the "Overspeed Margin" in the above VG diagram is the 10% margin between Vne and Vd as defined under FAR 23.1505?
Or is there some other margin above Vne in a VG diagram under FAR Part 23?
Or is there some other margin above Vne in a VG diagram under FAR Part 23?
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Speed as such is not a 'load'. 2.5g at 250kts and a given weight/CG is the same load on the structure as 2.5g at 150 kts and that weight/CG (those speeds chosen to eliminate compressibility effects). The loadings are defined at specific speeds so speed cannot really be also subject to a 1.5 factor.
Not sure what you have in mind here.
Basically that tells me it applies to an aircraft which is not experiencing changes in altitude and airspeed which would induce flutter at a speed lower than 1.15 Vd.
The following I posted above your last post, but you may miss it since I need to wait for mod approval....
In short, I am having a discussion with someone who feels the Loads defined in FAR 25.301, the factor of Safety in FAR 25.303, and the Strength and Deformation definitions in FAR 25.305 will provide a margin of safety for an aircraft which hypothetically would be accelerating through Vd+50 and pulling upward of 2-3.8g.
I contend no such margin exists nor is defined by Part 25.301, 303 and 305.
Thanks again for all your replies.
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They just don't build them like the used to
Read this NTSB report from the mid 70's regarding TWA flight 841. A 727 that fell from flight level 390. The old style flight recorder showed 6g and 470Kts at one point.
I wouldn't want to test that in an Airbus.
http://libraryonline.erau.edu/online...s/AAR81-08.pdf
I wouldn't want to test that in an Airbus.
http://libraryonline.erau.edu/online...s/AAR81-08.pdf
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Factor of Safety of 1.5
Bill,
This factor of safety is also used on the components and systems of the aircraft build-up.
Hydraulic pipes and cables are proof tested 1.5 times the normal operating value, wings are over bent to demine wing loading and fuselage pressure cycled such that when you fly your aircraft the whole 'component' has been well and truly proof tested to cover your normal flying plus.
Don't ask about the computers please......
This factor of safety is also used on the components and systems of the aircraft build-up.
Hydraulic pipes and cables are proof tested 1.5 times the normal operating value, wings are over bent to demine wing loading and fuselage pressure cycled such that when you fly your aircraft the whole 'component' has been well and truly proof tested to cover your normal flying plus.
Don't ask about the computers please......
Last edited by Tinwacker; 1st Jul 2014 at 12:47. Reason: Small ommission
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In short, I am having a discussion with someone who feels the Loads defined in FAR 25.301, the factor of Safety in FAR 25.303, and the Strength and Deformation definitions in FAR 25.305 will provide a margin of safety for an aircraft which hypothetically would be accelerating through Vd+50 and pulling upward of 2-3.8g.
I contend no such margin exists nor is defined by Part 25.301, 303 and 305.
I contend no such margin exists nor is defined by Part 25.301, 303 and 305.
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It would be a foolhardy man who took the margins in the regs and decided they were all available for use "on demand".
the aeroplane has required substantial rebuilds and has almost structurally failed in flight a number of times.
I know, its a FAR23 aircraft not FAR25 but there really are pilots out there who don't really understand all their knowledge.
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the aeroplane has required substantial rebuilds and has almost structurally failed in flight a number of times.
.. why are we not surprised ?
The fellow involved ought not to buy lottery tickets .. his luck has been used up several times over in his flying experiences.
.. why are we not surprised ?
The fellow involved ought not to buy lottery tickets .. his luck has been used up several times over in his flying experiences.
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Read this NTSB report from the mid 70's regarding TWA flight 841. A 727 that fell from flight level 390. The old style flight recorder showed 6g and 470Kts at one point.
The No. 7 leading edge slat on the right wing was missing. The slat tracks remained on the aircraft; the outboard track was twisted and bent rearward about midspan, and the inboard track was bent rearward near the aft end of the track. The slat actuator cylinder was broken about 1 1/2 inches forward of its trunnion; the aft portion of the cylinder remained attached to the wing. The forward end of the actuator cylinder, the actuator piston, and the piston rod were missing. The 5/16-inch bolts that attach the slat to its track were sheared, The inboard fairing-adjustment T-bolt was broken, and the threaded pqtion of the bolt and two adjusting nuts were missing...
The skin of the lower surface of the wing aft of the No. 7 slat actuator was scraped. An 8- to 10-inch portion of the outboard aileron balance tab was missing at the end of the scrape mark. The balance tab actuator lugs had separated, and the hinge support fitting between the lugs had sheared.
The right outboard aileron actuator hinge fitting bolt was broken. With the aileron in the locked-out position, there was free movement of 1 inch up and 3/32 inch down at the trailing edge of the aileron. The nut end of the bolt remained in the structure. A metallurgical examination of the bolt indicated that it had failed predominantly in fatigue.
The No. 10 flight spoiler panel, except for a portion containing the two inboard hinges, was missing. The right inboard trailing edge flap track attachment bolts were sheared and the carriage was damaged. The canoe-shaped fairing for the track was missing.
The No. 7 leading edge slat, which had broken into two pieces, and the outboard trailing edge flap track canoe-shaped fairing were found about 7 miles north of Saainaw. Michigan, at latitude 43'39'N and longitude 84%5'W. A large portion of the No. 10-spoiler panel was found about 3/4 mile south of these components. The forward portion of the No. 7 slat actuator cylinder, the actuator piston, and the piston rod were not found. The piston rod-end bearing remained attached to the slat; the rod had fractured in overload about 2 inches aft of the center of the bearing.
A metallurgical examination of the No. 7 slat inboard T-bolt indicated that3 the cross section of the bolt had fatigue fracture characteristics. There was considerable smearing of the fracture face.
Both main gear landing doors and their operating mechanisms were damaged3 extensively and a hydraulic line was ruptured. The sidebrace and actuator beam on the right gear were broken; the support beam for the left gear was intact. The uplock for the left gear was bent. The secondary wing skin panels above both actuator support beams were buckled upward.
The No. 4 flight spoiler was torn around its actuator attachment point....
The left outboard aileron balance tab hinge fitting was broken; in the locked-out position, there was no appreciable free movement of the aileron.
Many passenger oxygen masks were hanging from their overhead compartments. A passenger service unit was loosened from its moorings and an interior window was cracked.
The "A" hydraulic system reservoir contained 2 quarts of fluid. Following repair of the hydraulic line in the right wheel well and plugging of the No. 7 slat actuator lines, the reservoir was serviced and the flight controls and speed brakes were checked; they functioned properly. Except for the No. 7 leading edge slat, the leading edge slats and flaps, trailing edge flaps, and their indicator lights functioned properly on both the normal and alternate flap systems. The inboard could not be tested because of the damage to the right inboard trailing edge flap. The stall warning and overspeed warning systems functioned properly.
Slight tension-field wrinkles had formed in the fuselage skin fore and aft. - Source - http://www.airdisaster.com/reports/ntsb/AAR81-08.pdf
The skin of the lower surface of the wing aft of the No. 7 slat actuator was scraped. An 8- to 10-inch portion of the outboard aileron balance tab was missing at the end of the scrape mark. The balance tab actuator lugs had separated, and the hinge support fitting between the lugs had sheared.
The right outboard aileron actuator hinge fitting bolt was broken. With the aileron in the locked-out position, there was free movement of 1 inch up and 3/32 inch down at the trailing edge of the aileron. The nut end of the bolt remained in the structure. A metallurgical examination of the bolt indicated that it had failed predominantly in fatigue.
The No. 10 flight spoiler panel, except for a portion containing the two inboard hinges, was missing. The right inboard trailing edge flap track attachment bolts were sheared and the carriage was damaged. The canoe-shaped fairing for the track was missing.
The No. 7 leading edge slat, which had broken into two pieces, and the outboard trailing edge flap track canoe-shaped fairing were found about 7 miles north of Saainaw. Michigan, at latitude 43'39'N and longitude 84%5'W. A large portion of the No. 10-spoiler panel was found about 3/4 mile south of these components. The forward portion of the No. 7 slat actuator cylinder, the actuator piston, and the piston rod were not found. The piston rod-end bearing remained attached to the slat; the rod had fractured in overload about 2 inches aft of the center of the bearing.
A metallurgical examination of the No. 7 slat inboard T-bolt indicated that3 the cross section of the bolt had fatigue fracture characteristics. There was considerable smearing of the fracture face.
Both main gear landing doors and their operating mechanisms were damaged3 extensively and a hydraulic line was ruptured. The sidebrace and actuator beam on the right gear were broken; the support beam for the left gear was intact. The uplock for the left gear was bent. The secondary wing skin panels above both actuator support beams were buckled upward.
The No. 4 flight spoiler was torn around its actuator attachment point....
The left outboard aileron balance tab hinge fitting was broken; in the locked-out position, there was no appreciable free movement of the aileron.
Many passenger oxygen masks were hanging from their overhead compartments. A passenger service unit was loosened from its moorings and an interior window was cracked.
The "A" hydraulic system reservoir contained 2 quarts of fluid. Following repair of the hydraulic line in the right wheel well and plugging of the No. 7 slat actuator lines, the reservoir was serviced and the flight controls and speed brakes were checked; they functioned properly. Except for the No. 7 leading edge slat, the leading edge slats and flaps, trailing edge flaps, and their indicator lights functioned properly on both the normal and alternate flap systems. The inboard could not be tested because of the damage to the right inboard trailing edge flap. The stall warning and overspeed warning systems functioned properly.
Slight tension-field wrinkles had formed in the fuselage skin fore and aft. - Source - http://www.airdisaster.com/reports/ntsb/AAR81-08.pdf
I wouldn't want to test that in an Airbus.
The dive speed [Vd] is the absolute maximum speed above which the aircraft must not fly. Typically, to achieve this speed, the aircraft must enter a dive (steep descent), as the engines cannot produce sufficient thrust to overcome aerodynamic drag in level flight. At the dive speed, excessive aircraft vibrations develop which put the aircraft structural integrity at stake. Source - VD/MD | The Flying Engineer
Scroll down in the above source to see the video of flight testing certification out to the A380 Vd/Md.FAR 25.301, 303 and 305 do not pertain a 1.5 Factor of Safety in relation to speeds. In other words, if Vd were 400 knots, the 1.5 Factor of Safety defined under FAR 25.303 would not offer a 'margin of safety' in speed of (hypothetically) 420, or, 450, or 500... or 600 knots (1.5 x Vd).
Is there anyone here who disagrees?
Last edited by Bill Serger; 2nd Jul 2014 at 23:08.
§25.301 Loads.
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
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Read this NTSB report from the mid 70's regarding TWA flight 841. A 727 that fell from flight level 390. The old style flight recorder showed 6g and 470Kts at one point. I wouldn't want to test that in an Airbus.
Another extremely important aspect is, that the load is multiplied by a factor of 1.5, and not the material strength reduced by a factor of 1.5. What sounds equivalent at first, becomes relevant when we talk about stability (e.g. buckling) or internal loads of a deflected structure. An item loaded in bending and compression (e.g. members of an upper wing skin) sees more than 1.5 times the limit stress, when you load it to 1.5 times of limit load. The load stress relation becomes nonlinear if you take into account deformation, which occurs quite a bit on aircraft, as all of us have probably already seen.
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at lower weight more g is possible
.. for those without Volume's background ... the bits in the aircraft and bolted to the tin are designed for specific restraint factors and don't have the weight variation consideration .. ie the wing might not fall off but the pilot might have bits and pieces floating about his/her ears.
.. for those without Volume's background ... the bits in the aircraft and bolted to the tin are designed for specific restraint factors and don't have the weight variation consideration .. ie the wing might not fall off but the pilot might have bits and pieces floating about his/her ears.
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Loss of Safety factor
The 1.5 safety factor is also generally applied to military aircraft.
One popular light attack jet had wing tip lights fastened by 6 screws, 3 on top of the wing, and 3 on the bottom of the wing.
Aircrews and maintenance workers did not generally understand that flying with one screw missing then provided NO safety factor for that particular piece of equipment. Aircraft would often lose these lights on flights where the maximum allowed g was being approached.
One popular light attack jet had wing tip lights fastened by 6 screws, 3 on top of the wing, and 3 on the bottom of the wing.
Aircrews and maintenance workers did not generally understand that flying with one screw missing then provided NO safety factor for that particular piece of equipment. Aircraft would often lose these lights on flights where the maximum allowed g was being approached.
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....at lower weight more g is possible
At a lower weight, Va/Vra actually decreases. So, are you saying more G is possible at a decreased Vra due to a lower weight?
Or can one increase G loading at the same Va/Vra for a lower weight?