As a first post in a very interesting forum, I would like to ask 3rd party repair engineers what they use as an ultimate static stress (tensile) when designing “scab patches” on the fuselage skin. I think using Ftu leads to an over dimensioned repair when you consider that the fuselage is basically a thin walled tube in bending which fails in compression and leaves lots of opportunity to allow for local yielding in the upper shell. I think Fty is more appropriate. (reasons will come later)
Tek-Tips is the largest IT community on the Internet today!
Members share and learn making Tek-Tips Forums the best source of peer-reviewed technical information on the Internet!
-
Congratulations cowski on being selected by the Eng-Tips community for having the most helpful posts in the forums last week. Way to Go!
Ultimate stress in a pressurized fuselage 2
- Thread starter Bazzo
- Start date
I see that I am a little late to this post. However, I just recently joined and I came across this one during my reading.
I think that one important thing that is forgotten is the FAR's 25.301, 303, 305, 307. FAR 25.303 states that a 1.5 factor must be applied to the limit load during design. Even though the normal operating or limit design load maybe less than the actual Fty. The 1.5(limit load) may be greater than the Fty. Therefore, since you do not have the actual limit load or normal operting load of the skin panel, you must use 1.5Fty or Ftu (conserviative)in order to meet this FAR's. As a DER, showing compliance with the FAR's or CFR's is req'd.
I think that one important thing that is forgotten is the FAR's 25.301, 303, 305, 307. FAR 25.303 states that a 1.5 factor must be applied to the limit load during design. Even though the normal operating or limit design load maybe less than the actual Fty. The 1.5(limit load) may be greater than the Fty. Therefore, since you do not have the actual limit load or normal operting load of the skin panel, you must use 1.5Fty or Ftu (conserviative)in order to meet this FAR's. As a DER, showing compliance with the FAR's or CFR's is req'd.
dave081069
New member
In their Part III Structures course Boeing recommend sizing a skin repair to the lowest of:
1. Ftu
2. The load capability of the applicable SRM repair, or
3. the load capability of adjacent skin production joints.
Obviously all other rules including doubler thickness, stepping/tapering and appropriate fasteners apply.
1. Ftu
2. The load capability of the applicable SRM repair, or
3. the load capability of adjacent skin production joints.
Obviously all other rules including doubler thickness, stepping/tapering and appropriate fasteners apply.
I agree with dave0806, that without having the actual running loads, one of those three ways he has listed is the way to determine the ultimate load associated with the panel. You can show compliance with the FAR's using one of those three methods. Just be sure if you use no 2 - the load capability of the applicable SRM repair that the repair is actually applicable to your skin panel. As a DER I have reviewed Engineer's skin repairs that have referenced SRM repairs that stated in the applicability block that it was applicable to every skin panel but the one they were repairing. Not saying that the repair wasn't adequate, they just couldn't use the SRM as their basis of approval.
On a different note -I must say that I am a newbie to this forum, but really enjoy it. I have noticed that I have to read my own posts a few times because what I am thinking doesn't always come out on the keyboard the correct way!! Please let me know if any of my posts don't make complete sense.
On a different note -I must say that I am a newbie to this forum, but really enjoy it. I have noticed that I have to read my own posts a few times because what I am thinking doesn't always come out on the keyboard the correct way!! Please let me know if any of my posts don't make complete sense.
I am glad this thread is revived.
Dave081069:
-What do you mean by point #2? Do you just mean following the SRM?
-There's a dichotomy between the first two points and the last point. If you look at a skin joint on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows!! What gives?
737eng:
Do structures DER's have access to OEM loads? I wonder what basis you use to sign those 8110's. Do you simply ensure that the repair meets Ftu, and that it follows general durability guidelines? I am curious as to whether the DER's job leans more towards verifying calculations, strength, etc. (i.e. engineering) versus ensuring actual FAR compliance (i.e. regulatory paperwork)
Does bumping up Ftu by 15% (the fitting factor) apply to all structural joints?
Cheers,
Alex
Dave081069:
-What do you mean by point #2? Do you just mean following the SRM?
-There's a dichotomy between the first two points and the last point. If you look at a skin joint on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows!! What gives?
737eng:
Do structures DER's have access to OEM loads? I wonder what basis you use to sign those 8110's. Do you simply ensure that the repair meets Ftu, and that it follows general durability guidelines? I am curious as to whether the DER's job leans more towards verifying calculations, strength, etc. (i.e. engineering) versus ensuring actual FAR compliance (i.e. regulatory paperwork)
Does bumping up Ftu by 15% (the fitting factor) apply to all structural joints?
Cheers,
Alex
- Thread starter
- #25
Interesting to hear that you are all happy using Ftu. If you assume your structure is working at Ftu you cannot put any holes in it for fasteners etc. It breaks. As regards the Boeing course, the Airbus course says it’s ok to use Fty as the applied ultimate stress for skin repairs, i.e. the stress in the structure resulting from a loading which is equal to 1.5 x the limit load.
dave081069
New member
Koopas
#2 is really about using the SRM repair as a method of determining the loads required for a modifed repair i.e. a repair in a location or configuration where a standard SRM repair cannot be accomplished.
#2 is really about using the SRM repair as a method of determining the loads required for a modifed repair i.e. a repair in a location or configuration where a standard SRM repair cannot be accomplished.
Bazzo:
When you say that your doubler will break if you assume the structure works at Ftu due to the fastener holes, wouldn't going the standard "one gage up" account for the reduction in net tension area?
In one of your replies above, you mention three points to design to Fty instead of Ftu:
Quoted:
"1. The reason in the first post, local yielding in the thin walled tube (fuselage) will not permit the tension stress to go beyond Fty, it will always find a way to include more effective tension area.
2. Modern day jet liners (including those using Glare) are designed to a (undisturbed area) fatigue life of 90 to 110 N/mm² fatigue stress converting this to an ultimate static stress (Nz = 2,5 and 1,5 delta P) you obtain a stress round about yield.
3. (maybe a little weak but) Linear FEM models do not obviously include the inelastic behaviour of the structure, reducing the ultimate elastic stress (Ftu value) using Ramberg-Osgood you get a value under Fty when considering 2024 T3 properties."
Could you please enlighten me a little and digress a bit on those points? I am a little lost.
Thanks,
Alex
When you say that your doubler will break if you assume the structure works at Ftu due to the fastener holes, wouldn't going the standard "one gage up" account for the reduction in net tension area?
In one of your replies above, you mention three points to design to Fty instead of Ftu:
Quoted:
"1. The reason in the first post, local yielding in the thin walled tube (fuselage) will not permit the tension stress to go beyond Fty, it will always find a way to include more effective tension area.
2. Modern day jet liners (including those using Glare) are designed to a (undisturbed area) fatigue life of 90 to 110 N/mm² fatigue stress converting this to an ultimate static stress (Nz = 2,5 and 1,5 delta P) you obtain a stress round about yield.
3. (maybe a little weak but) Linear FEM models do not obviously include the inelastic behaviour of the structure, reducing the ultimate elastic stress (Ftu value) using Ramberg-Osgood you get a value under Fty when considering 2024 T3 properties."
Could you please enlighten me a little and digress a bit on those points? I am a little lost.
Thanks,
Alex
Dave081069:
Thanks for your reply. In my measly 15 months in the biz, it's never come up that I've had to use the SRM to attempt to back out the loads. Usually, if there's an SRM repair for that area, we use the SRM repair. If not, we figure out the ultimate loads and size the doubler/fasteners accordingly. I am wondering if you could share your experiences on the issue.
Oh yeah, do you also have any comments on my second point in my last post?
Cheers,
Alex
Thanks for your reply. In my measly 15 months in the biz, it's never come up that I've had to use the SRM to attempt to back out the loads. Usually, if there's an SRM repair for that area, we use the SRM repair. If not, we figure out the ultimate loads and size the doubler/fasteners accordingly. I am wondering if you could share your experiences on the issue.
Oh yeah, do you also have any comments on my second point in my last post?
Cheers,
Alex
- Thread starter
- #29
Alex.
I slight misunderstanding, I wasn’t talking about the doubler, I meant the original structure, which you obviously can’t change, unless you’re thinking of hot-bonding a doubler to it.
As to the reasons:
Progressively bend the fuselage so as to develop tension in the upper shell and compression in the lower shell. The upper shell stresses will increase linearly to the Fty value in the crown area. When they reach this value the crown area will in effect increase due to local yielding but the stress will remain constant. This area will con-tinue to increase until the fuselage fails in compression in the lower shell. The fuse-lage is simply not massive enough to go plastic in the bending case. Imagine bending a coke can, when you bend it, it always fails in compression.
You could try optimizing the fuselage (i.e. making it thinner in the upper shell area), which may help a little, but then the fatigue engineer shows up and says with a fa-tigue stress greater than e.g. 110 N/mm² he is not able to show a usable life. This fatigue stress is with an Nz = 1 and at 1 Delta P (i.e. internal pressure differential of 1 which is about 9 psi). The ultimate case has an Nz of 2,5 and 1,5 Delta P (FAR 25). So you could factor this 110 N/mm² up and still be in the area of Fty. In addition there are other cases where the upper shell is also in compression (e.g. emergency brak-ing case) so you have to design a bit of stability, so you can’t get too thin.
The third reason assumed that the manufacturer uses a linear FEM calculation (I did say it was weak). If you reduce a linear stress, which is equal to the Ftu value of 2024, T3 following Ramberg-Osgood principles, you obtain a value below the actual Fty value. The manufacturers FEM program will not allow him to go beyond the Ftu value as he would have an MoS less than zero.
The above probably produces more questions than it answers but I hope it’s a bit of help in trying to explain my point.
I slight misunderstanding, I wasn’t talking about the doubler, I meant the original structure, which you obviously can’t change, unless you’re thinking of hot-bonding a doubler to it.
As to the reasons:
Progressively bend the fuselage so as to develop tension in the upper shell and compression in the lower shell. The upper shell stresses will increase linearly to the Fty value in the crown area. When they reach this value the crown area will in effect increase due to local yielding but the stress will remain constant. This area will con-tinue to increase until the fuselage fails in compression in the lower shell. The fuse-lage is simply not massive enough to go plastic in the bending case. Imagine bending a coke can, when you bend it, it always fails in compression.
You could try optimizing the fuselage (i.e. making it thinner in the upper shell area), which may help a little, but then the fatigue engineer shows up and says with a fa-tigue stress greater than e.g. 110 N/mm² he is not able to show a usable life. This fatigue stress is with an Nz = 1 and at 1 Delta P (i.e. internal pressure differential of 1 which is about 9 psi). The ultimate case has an Nz of 2,5 and 1,5 Delta P (FAR 25). So you could factor this 110 N/mm² up and still be in the area of Fty. In addition there are other cases where the upper shell is also in compression (e.g. emergency brak-ing case) so you have to design a bit of stability, so you can’t get too thin.
The third reason assumed that the manufacturer uses a linear FEM calculation (I did say it was weak). If you reduce a linear stress, which is equal to the Ftu value of 2024, T3 following Ramberg-Osgood principles, you obtain a value below the actual Fty value. The manufacturers FEM program will not allow him to go beyond the Ftu value as he would have an MoS less than zero.
The above probably produces more questions than it answers but I hope it’s a bit of help in trying to explain my point.
Quote:
Do structures DER's have access to OEM loads? I wonder what basis you use to sign those 8110's. Do you simply ensure that the repair meets Ftu, and that it follows general durability guidelines? I am curious as to whether the DER's job leans more towards verifying calculations, strength, etc. (i.e. engineering) versus ensuring actual FAR compliance (i.e. regulatory paperwork)
Does bumping up Ftu by 15% (the fitting factor) apply to all structural joints?
Alex,
Only DER's at OEM's have access to this proprietary information. DER's in the field typically do not have load data. Even though I chose the username 737eng, I do not work for the big B, I only chose that name since I have come to enjoy the 737 over the years and it has kept me employed.
The DER finds compliance with the FAR's. In order to show compliance a structural analysis typically has to be accomplished or reviewed. (other data and/or analysis may have to be accomplished as well, i.e. DT, Burn tests, electrical load, emi, etc....however I am a structures DER without DT so I am only showing compliance to FAR's that deal with static strength). A DER has a mixture of both regulatory paperwork and actual engineering.
By the way, I still call them FAR's, however, I believe the correct terminology is Code of Federal Regulation, Title 14 or CFR 14.
When you state bumping up FTU by 15% are you talking about the 1.15 fitting factor per FAR 25.625? If so, the 1.15 fitting factor only needs to be utilized when you are not utilizing tested data for your allowables.
Bazzo,
I feel that the belly and the crown are definetly designed differently to account for the different loading. Typically the belly skins are thicker, plus, the stringers are designed to take the compressive loading.
I understand your statement that the structure never sees FTU, otherwise we couldn't add any fastener holes. However, still feel the key is that since you do not know the actual limit load in the panel, which can be anywhere from (-) all the way up to Fty. You cannot verify that you meet the 1.5(limit load) req'd by the FAR's. Let me know your thoughts on this issue.
Do structures DER's have access to OEM loads? I wonder what basis you use to sign those 8110's. Do you simply ensure that the repair meets Ftu, and that it follows general durability guidelines? I am curious as to whether the DER's job leans more towards verifying calculations, strength, etc. (i.e. engineering) versus ensuring actual FAR compliance (i.e. regulatory paperwork)
Does bumping up Ftu by 15% (the fitting factor) apply to all structural joints?
Alex,
Only DER's at OEM's have access to this proprietary information. DER's in the field typically do not have load data. Even though I chose the username 737eng, I do not work for the big B, I only chose that name since I have come to enjoy the 737 over the years and it has kept me employed.
The DER finds compliance with the FAR's. In order to show compliance a structural analysis typically has to be accomplished or reviewed. (other data and/or analysis may have to be accomplished as well, i.e. DT, Burn tests, electrical load, emi, etc....however I am a structures DER without DT so I am only showing compliance to FAR's that deal with static strength). A DER has a mixture of both regulatory paperwork and actual engineering.
By the way, I still call them FAR's, however, I believe the correct terminology is Code of Federal Regulation, Title 14 or CFR 14.
When you state bumping up FTU by 15% are you talking about the 1.15 fitting factor per FAR 25.625? If so, the 1.15 fitting factor only needs to be utilized when you are not utilizing tested data for your allowables.
Bazzo,
I feel that the belly and the crown are definetly designed differently to account for the different loading. Typically the belly skins are thicker, plus, the stringers are designed to take the compressive loading.
I understand your statement that the structure never sees FTU, otherwise we couldn't add any fastener holes. However, still feel the key is that since you do not know the actual limit load in the panel, which can be anywhere from (-) all the way up to Fty. You cannot verify that you meet the 1.5(limit load) req'd by the FAR's. Let me know your thoughts on this issue.
737eng:
If I read you correctly, you're only checking for static strength. Some additional questions:
1. Do you base static strength on Ftu ONLY?
2. Do you ever do a damage tolerance analysis beyond the static strength analysis? Is the static analysis then forwarded to another DER who performs a DT analysis? It would appear that static strength alone is insufficient to show the acceptability of a repair.
3. Are all your static calculations hand-based? If not, what computer programs do you use? Do you use FEA?
4. How does one become a structures DER, beyond the application to the FAA? Is any particular experience or educational level required?
5. Yes, the 15% fitting factor you're describing is what I was talking about. When you write "tested allowables", do you mean that the FAR's require a 15% increase in Ftu if I don't obtain the joint allowables from the SRM or Mil-Handbook 5? Douglas training courses add the 15% regardless. Boeing training courses never do. That said, thanks for listing the FAR number, I'll look it up.
6. When you can't find a joint allowable for a certain fastener/sheet material combination under single shear, do you find it acceptable to use Fbru * d * t (even though Fbru is based on double shear, with less eccentricity?) to determine the joint allowable, provided that the double-shear strength of the fastener is higher?
Thanks for providing some insight.
Alex
If I read you correctly, you're only checking for static strength. Some additional questions:
1. Do you base static strength on Ftu ONLY?
2. Do you ever do a damage tolerance analysis beyond the static strength analysis? Is the static analysis then forwarded to another DER who performs a DT analysis? It would appear that static strength alone is insufficient to show the acceptability of a repair.
3. Are all your static calculations hand-based? If not, what computer programs do you use? Do you use FEA?
4. How does one become a structures DER, beyond the application to the FAA? Is any particular experience or educational level required?
5. Yes, the 15% fitting factor you're describing is what I was talking about. When you write "tested allowables", do you mean that the FAR's require a 15% increase in Ftu if I don't obtain the joint allowables from the SRM or Mil-Handbook 5? Douglas training courses add the 15% regardless. Boeing training courses never do. That said, thanks for listing the FAR number, I'll look it up.
6. When you can't find a joint allowable for a certain fastener/sheet material combination under single shear, do you find it acceptable to use Fbru * d * t (even though Fbru is based on double shear, with less eccentricity?) to determine the joint allowable, provided that the double-shear strength of the fastener is higher?
Thanks for providing some insight.
Alex
Alex,
1. Do you base static strength on Ftu ONLY?
No, I base the static strength on the ultimate load carrying capability of the structural member (i.e. skin panel per this post). As can be determined from reading all the above posts, there is more than one way to determine the actual ultimate "running" load of the panel. If I cannot determine the actual ultimate load carrying capability or if I am pressed for time, I will then resort back to using Ftu, knowing this is extremely conservative.
2. Do you ever do a damage tolerance analysis beyond the static strength analysis? Is the static analysis then forwarded to another DER who performs a DT analysis? It would appear that static strength alone is insufficient to show the acceptability of a repair.
Currently, I typically am involved with pre-ammendment 45 aircraft and therefore do not have to show compliance with 25-571 (b) {don't quote me on the actual FAR). I typically just have to meet the "fail-safe" requirements of 25.571 since that is what the aircraft was certified to. I can show that I meet fail-safe by either evaluating the original design, or the SRM. This will change in the future, since Congress and thus the FAA are driving us into basically making all a/c damage tolerant. However, when I do get the opportunity to work on a DT aircraft, I cannot provide the DT approval. Therefore, I either utilize the guidance of a pre-existing repair (i.e. SRM) or I limit my approval to one-year following the guidelines of AC 25.1529 and then submit to obtain the DT approval (typically the OEM).
3. Are all your static calculations hand-based? If not, what computer programs do you use? Do you use FEA?
Most of my calculations utilize classic "hand-based" structural analysis. I do utilize computer programs, such as ACAD Mass Properties, Virtual DER, excel (build a spreadsheet to do crippling analysis, etc). I do not use FEA software.
4. How does one become a structures DER, beyond the application to the FAA? Is any particular experience or educational level required?
The application and requirements are on the FAA website. In general you basically need 8 yrs experience in the area you are applying and have to prove that you are familiar with the FAR's and dealing with the FAA.
5. Yes, the 15% fitting factor you're describing is what I was talking about. When you write "tested allowables", do you mean that the FAR's require a 15% increase in Ftu if I don't obtain the joint allowables from the SRM or Mil-Handbook 5? Douglas training courses add the 15% regardless. Boeing training courses never do. That said, thanks for listing the FAR number, I'll look it up.
Look up the FAR, it provides the guidance you need.
6. When you can't find a joint allowable for a certain fastener/sheet material combination under single shear, do you find it acceptable to use Fbru * d * t (even though Fbru is based on double shear, with less eccentricity?) to determine the joint allowable, provided that the double-shear strength of the fastener is higher?
I do utilize Fbru * d * t as long as it is a protruding head fastener or the buck tail is against the sheet. I do not feel that Fbru * d * t is acceptable for countersunk fasteners.
1. Do you base static strength on Ftu ONLY?
No, I base the static strength on the ultimate load carrying capability of the structural member (i.e. skin panel per this post). As can be determined from reading all the above posts, there is more than one way to determine the actual ultimate "running" load of the panel. If I cannot determine the actual ultimate load carrying capability or if I am pressed for time, I will then resort back to using Ftu, knowing this is extremely conservative.
2. Do you ever do a damage tolerance analysis beyond the static strength analysis? Is the static analysis then forwarded to another DER who performs a DT analysis? It would appear that static strength alone is insufficient to show the acceptability of a repair.
Currently, I typically am involved with pre-ammendment 45 aircraft and therefore do not have to show compliance with 25-571 (b) {don't quote me on the actual FAR). I typically just have to meet the "fail-safe" requirements of 25.571 since that is what the aircraft was certified to. I can show that I meet fail-safe by either evaluating the original design, or the SRM. This will change in the future, since Congress and thus the FAA are driving us into basically making all a/c damage tolerant. However, when I do get the opportunity to work on a DT aircraft, I cannot provide the DT approval. Therefore, I either utilize the guidance of a pre-existing repair (i.e. SRM) or I limit my approval to one-year following the guidelines of AC 25.1529 and then submit to obtain the DT approval (typically the OEM).
3. Are all your static calculations hand-based? If not, what computer programs do you use? Do you use FEA?
Most of my calculations utilize classic "hand-based" structural analysis. I do utilize computer programs, such as ACAD Mass Properties, Virtual DER, excel (build a spreadsheet to do crippling analysis, etc). I do not use FEA software.
4. How does one become a structures DER, beyond the application to the FAA? Is any particular experience or educational level required?
The application and requirements are on the FAA website. In general you basically need 8 yrs experience in the area you are applying and have to prove that you are familiar with the FAR's and dealing with the FAA.
5. Yes, the 15% fitting factor you're describing is what I was talking about. When you write "tested allowables", do you mean that the FAR's require a 15% increase in Ftu if I don't obtain the joint allowables from the SRM or Mil-Handbook 5? Douglas training courses add the 15% regardless. Boeing training courses never do. That said, thanks for listing the FAR number, I'll look it up.
Look up the FAR, it provides the guidance you need.
6. When you can't find a joint allowable for a certain fastener/sheet material combination under single shear, do you find it acceptable to use Fbru * d * t (even though Fbru is based on double shear, with less eccentricity?) to determine the joint allowable, provided that the double-shear strength of the fastener is higher?
I do utilize Fbru * d * t as long as it is a protruding head fastener or the buck tail is against the sheet. I do not feel that Fbru * d * t is acceptable for countersunk fasteners.
Hello 737eng:
1. Referring to your reply's first answer, and the determination of load capability in a skin panel, I'll repost my comment made to another poster: If you look at a skin production splice on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5, sometimes 4.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows. What gives? How does one justify the variation from two to six rows?
2. Don't you think that using Fbru*d*t for a protruding head fastener over-estimates the single-shear strength of a joint, due to the added joint bending compared to a double-shear joint?
3. How do you compute the bearing strength of a countersunk fastener in a single-shear joint?
Regards,
Alex
1. Referring to your reply's first answer, and the determination of load capability in a skin panel, I'll repost my comment made to another poster: If you look at a skin production splice on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5, sometimes 4.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows. What gives? How does one justify the variation from two to six rows?
2. Don't you think that using Fbru*d*t for a protruding head fastener over-estimates the single-shear strength of a joint, due to the added joint bending compared to a double-shear joint?
3. How do you compute the bearing strength of a countersunk fastener in a single-shear joint?
Regards,
Alex
Alex,
1. Referring to your reply's first answer, and the determination of load capability in a skin panel, I'll repost my comment made to another poster: If you look at a skin production splice on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5, sometimes 4.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows. What gives? How does one justify the variation from two to six rows?
There could be multiple reasons for this, which I do not have time to look into right now:
I am not that familiar with the B717, are the skins chem-milled? If so, the thickness of the skin (or added doublers/finger doublers if not chem-milled) is most likely greater than the skin thickness in the pocket (between stringer and frames). Thus due to the reduced skin thickness the fastener bearing capability is reduced resulting in the need for more fastener rows.
2. Don't you think that using Fbru*d*t for a protruding head fastener over-estimates the single-shear strength of a joint, due to the added joint bending compared to a double-shear joint?
You should always use tested data, if not, the 1.15 fitting factor per FAR 25.625 should be utilized. Therefore, if you calculate your bearing Fbru*d*t you will need to utilize the 1.15 which will account for transitional failure modes.
3. How do you compute the bearing strength of a countersunk fastener in a single-shear joint?
You have to use tested data. I do not know how to calculate it. If you or someone else know how, I would be delighted to find out.
1. Referring to your reply's first answer, and the determination of load capability in a skin panel, I'll repost my comment made to another poster: If you look at a skin production splice on the 717, you'll see two rows of fasteners. The skin repair section in Chapter 53 mandates a minimum of 3.5, sometimes 4.5 rows (with finger doublers) for permanent repairs. If you were to use Ftu, you'd probably end up with 5 or 6 rows. What gives? How does one justify the variation from two to six rows?
There could be multiple reasons for this, which I do not have time to look into right now:
I am not that familiar with the B717, are the skins chem-milled? If so, the thickness of the skin (or added doublers/finger doublers if not chem-milled) is most likely greater than the skin thickness in the pocket (between stringer and frames). Thus due to the reduced skin thickness the fastener bearing capability is reduced resulting in the need for more fastener rows.
2. Don't you think that using Fbru*d*t for a protruding head fastener over-estimates the single-shear strength of a joint, due to the added joint bending compared to a double-shear joint?
You should always use tested data, if not, the 1.15 fitting factor per FAR 25.625 should be utilized. Therefore, if you calculate your bearing Fbru*d*t you will need to utilize the 1.15 which will account for transitional failure modes.
3. How do you compute the bearing strength of a countersunk fastener in a single-shear joint?
You have to use tested data. I do not know how to calculate it. If you or someone else know how, I would be delighted to find out.
-
1
- #35
Hey all,
This is a great forum, lots of good discussion. I'll wade in with my 2 cents worth.
As for designing repairs, the use of MATERIAL FTU alone is overly conservative. One approach is to look at a nearby section, such as a skin joint with rivets connecting to a stringer. At this location, the maximum hoop load is based on the NET AREA * FTU. That is the maximum load in that skin panel/bay. If it is a joint, the lesser of the fastener transfer capability or the net section ULTIMATE strength should be used.
Now it was said that Airbus lets one use Fty. That might be a design criteria they used from the begining, and that method should not be transfered to other aircraft. Boeing and other manufacturers all have their own criteria that must be met during design.
later,
jetmaker
This is a great forum, lots of good discussion. I'll wade in with my 2 cents worth.
As for designing repairs, the use of MATERIAL FTU alone is overly conservative. One approach is to look at a nearby section, such as a skin joint with rivets connecting to a stringer. At this location, the maximum hoop load is based on the NET AREA * FTU. That is the maximum load in that skin panel/bay. If it is a joint, the lesser of the fastener transfer capability or the net section ULTIMATE strength should be used.
Now it was said that Airbus lets one use Fty. That might be a design criteria they used from the begining, and that method should not be transfered to other aircraft. Boeing and other manufacturers all have their own criteria that must be met during design.
later,
jetmaker
Hi 737eng:
1. No, the areas of skins I've delt with (below the cusp line) on the 717 are not chem-milled unlike the Boeing models. However, you are correct in that there are finger doublers at factory skin splice locations thus increasing the local thickness at these locations. Consequently, I agree that the skin thickness in the "pocket", away from factory skin splices, would be relatively thinner. However, why would you size a repair doubler based on the thicker skin thickness at the factory skin splice, when the local thickness to be repaired is relatively thinner?
2. I understand that one must use tested data, when available. However, often enough, you find yourself with fastener/sheet material combinations not found in MIL-HNBK 5 or the SRM. I'd be interested in your comments regarding the overestimation of using the bearing formula Fbru*d*t for single-shear joints due to the added bending, and how you'd go about for correcting for that overestimation.
Hi Jetmaker:
Are you saying that the maximum panel load carrying capability in the HOOP direction is based on the NET AREA (gross area minus fastener areas) of a skin/stringer joint (not necessarily a skin splice location)? What about the maximum panel load carrying capability in the LONGITUDINAL direction?
You then mention "If it is a joint, the lesser of the fastener transfer capability or the net section ULTIMATE strength should be used." I am little confused as to the difference between this and my previous paragraph above when determining the max. load capability of a skin panel (I think it boils down to terminology confusion or rather terminology ignorance on my part
. Could you please clarify the nuance?
Thanks,
A curious Alex
1. No, the areas of skins I've delt with (below the cusp line) on the 717 are not chem-milled unlike the Boeing models. However, you are correct in that there are finger doublers at factory skin splice locations thus increasing the local thickness at these locations. Consequently, I agree that the skin thickness in the "pocket", away from factory skin splices, would be relatively thinner. However, why would you size a repair doubler based on the thicker skin thickness at the factory skin splice, when the local thickness to be repaired is relatively thinner?
2. I understand that one must use tested data, when available. However, often enough, you find yourself with fastener/sheet material combinations not found in MIL-HNBK 5 or the SRM. I'd be interested in your comments regarding the overestimation of using the bearing formula Fbru*d*t for single-shear joints due to the added bending, and how you'd go about for correcting for that overestimation.
Hi Jetmaker:
Are you saying that the maximum panel load carrying capability in the HOOP direction is based on the NET AREA (gross area minus fastener areas) of a skin/stringer joint (not necessarily a skin splice location)? What about the maximum panel load carrying capability in the LONGITUDINAL direction?
You then mention "If it is a joint, the lesser of the fastener transfer capability or the net section ULTIMATE strength should be used." I am little confused as to the difference between this and my previous paragraph above when determining the max. load capability of a skin panel (I think it boils down to terminology confusion or rather terminology ignorance on my part
. Could you please clarify the nuance?Thanks,
A curious Alex
koopas,
Yep.. you have it correct. The maximum panel capability is based on NET AREA = Gross Area - Hole Out Area. The same applies for the longitudinal axis. But note, this is not necessarily at stringer locations. This could be at a cutout, a row of rivets used in a doubler, or simply the gross area of a chem-milled pocket. Use whichever is less.
Let me try from a different approach with regards to the joint. If you have a 2 row lap joint, calculate the bearing/shear capability of each fastener. Then multiply the total fasteners in the lap by this number. This is the joint strength based on fastener capability. Next, check the NET AREA strength at the first row. Whichever is less is the joint strength capability which governs your design loads.
Hope this helped.
jetmaker
Yep.. you have it correct. The maximum panel capability is based on NET AREA = Gross Area - Hole Out Area. The same applies for the longitudinal axis. But note, this is not necessarily at stringer locations. This could be at a cutout, a row of rivets used in a doubler, or simply the gross area of a chem-milled pocket. Use whichever is less.
Let me try from a different approach with regards to the joint. If you have a 2 row lap joint, calculate the bearing/shear capability of each fastener. Then multiply the total fasteners in the lap by this number. This is the joint strength based on fastener capability. Next, check the NET AREA strength at the first row. Whichever is less is the joint strength capability which governs your design loads.
Hope this helped.
jetmaker
Hello Jetmaker,
Let me see if I've got this straight...
1. Essentially, you're saying that the max. NET AREA (i.e. material-based) load carrying capability of a skin panel is the smallest NET cross sectional area (Net Area_min) at ANY LOCATION in that panel (whether it's a skin splice location, a stringer to skin attachment location, etc.) times Ftu? Choosing the smallest net cross sectional area gives the maximum load that can be carried by the MATERIAL at any location.
2. Then, you compare it to the min. load carrying capability by the fasteners at any joint location on that panel (If you have two rows of 50 each -6 rivets at a butt joint along the panel, the max load carrying capability by the fasteners in that skin panel is 100 * Pall of a single -6 rivet in that sheet material/thickness). Choosing the minimum load carrying capability at any joint location on the panel gives the maximum load that can be carried by the FASTENERS at any location on the panel.
The lesser of the two (Net area_min*Ftu vs. min. fastener load carrying capability) is the strength of panel? In other words, the lesser of points 1 and 2 gives the maximum load carrying capability of the panel?
Did that sound right?
Cheers,
Alex
Let me see if I've got this straight...
1. Essentially, you're saying that the max. NET AREA (i.e. material-based) load carrying capability of a skin panel is the smallest NET cross sectional area (Net Area_min) at ANY LOCATION in that panel (whether it's a skin splice location, a stringer to skin attachment location, etc.) times Ftu? Choosing the smallest net cross sectional area gives the maximum load that can be carried by the MATERIAL at any location.
2. Then, you compare it to the min. load carrying capability by the fasteners at any joint location on that panel (If you have two rows of 50 each -6 rivets at a butt joint along the panel, the max load carrying capability by the fasteners in that skin panel is 100 * Pall of a single -6 rivet in that sheet material/thickness). Choosing the minimum load carrying capability at any joint location on the panel gives the maximum load that can be carried by the FASTENERS at any location on the panel.
The lesser of the two (Net area_min*Ftu vs. min. fastener load carrying capability) is the strength of panel? In other words, the lesser of points 1 and 2 gives the maximum load carrying capability of the panel?
Did that sound right?
Cheers,
Alex
koopas,
Sounds like you have a good understanding of the approach I use. However, just to clarify a few things. 1) the definition of panel in this case is a localized region of the skin, defined as a small region between frames and stringers. You can not use a lap joint capability to analyze a portion of the skin 5 stringers above the lap. 2) the strength of each panel must be calculated in each direction independantly. So, do not use a lap joint to calculate the tensile capability in the hoop and longitudinal direction. Make sure you choose wisely.
Regards,
jetmaker
Sounds like you have a good understanding of the approach I use. However, just to clarify a few things. 1) the definition of panel in this case is a localized region of the skin, defined as a small region between frames and stringers. You can not use a lap joint capability to analyze a portion of the skin 5 stringers above the lap. 2) the strength of each panel must be calculated in each direction independantly. So, do not use a lap joint to calculate the tensile capability in the hoop and longitudinal direction. Make sure you choose wisely.
Regards,
jetmaker
Greetings jetmaker,
Questions are numbered 1 to 6.
1. So, you only use a lap joint to calculate the tensile capability in the same direction as the lap joint? (i.e. if the lap joint runs fore-aft, the lap joint will give you the hoop tensile capability only?) For the longitudinal tensile capability, you might use a factory skin splice butt joint or even just Ftu is no splice joints are nearby? What about using a frame's net area (net area of skin minus fasteners' area common to skin & frame) or total fastener capability (Pall of skin/frame single shear joint), whichever is lower, to get the longitudinal tensile capability?
2. Could you digress a bit into "You can not use a lap joint capability to analyze a portion of the skin 5 stringers above the lap"? Say you'd want to find out the tensile capability of the that local area five stringers above the lap joint (considering the lap joint runs fore-aft):
You'd have to use the lesser of the net area*Ftu AT the stringer along that panel's length OR the total fastener capability (Pall of skin/stringer single shear joint) along that panel's length AT the stringer? In both cases, the stringer is the closest attaching structure to the skin near the local area of skin to be repaired. Is that correct?
3. However, if the "local" area to be analyzed was right above the lap, it'd be okay to use the lesser of the lap joint's net area*Ftu OR total fastener capability (Pall of the lap joint) to obtain the max. tensile strength of that local area in the hoop direction?
4. It would appear as though the hoop tensile capability near a lap joint would be higher than near a stringer since you might find higher skin thicknesses around lap joints, as well as a minimum of two rows of fasteners at the lap joint, versus a thinner skin and only one row of fasteners through the hat section of the stringer. Any comments?
5. Last point may sound trivial but I just want to make sure: Say you're above the lap joint as described above. So you use the lap joint as the reference for obtaining your net area*Ftu and total fastener load (you take the lesser of the two). Obviously, the lap joint runs for a good length along the skin panel's length (say 100 inches). Now, your cutout is only 10 inches in length. You'd only use 10 inches of net area for your material capability, right? Likewise, when calculating the total fastener capability, you'd only multiply a single fastener's joint allowable strength (Pall) times the the number of fasteners contained above the cutout within a 10-in length segment, right?
6. Say you need to cutout part of a beam's web. Instead of using Ftu, would it be acceptable to use Fsu (much lower than Ftu) for doubler sizing since the loading in the web is primarily shear?
Hope my questions make sense. It's late. Looking forward to reading your answers.
Alex
Questions are numbered 1 to 6.
1. So, you only use a lap joint to calculate the tensile capability in the same direction as the lap joint? (i.e. if the lap joint runs fore-aft, the lap joint will give you the hoop tensile capability only?) For the longitudinal tensile capability, you might use a factory skin splice butt joint or even just Ftu is no splice joints are nearby? What about using a frame's net area (net area of skin minus fasteners' area common to skin & frame) or total fastener capability (Pall of skin/frame single shear joint), whichever is lower, to get the longitudinal tensile capability?
2. Could you digress a bit into "You can not use a lap joint capability to analyze a portion of the skin 5 stringers above the lap"? Say you'd want to find out the tensile capability of the that local area five stringers above the lap joint (considering the lap joint runs fore-aft):
You'd have to use the lesser of the net area*Ftu AT the stringer along that panel's length OR the total fastener capability (Pall of skin/stringer single shear joint) along that panel's length AT the stringer? In both cases, the stringer is the closest attaching structure to the skin near the local area of skin to be repaired. Is that correct?
3. However, if the "local" area to be analyzed was right above the lap, it'd be okay to use the lesser of the lap joint's net area*Ftu OR total fastener capability (Pall of the lap joint) to obtain the max. tensile strength of that local area in the hoop direction?
4. It would appear as though the hoop tensile capability near a lap joint would be higher than near a stringer since you might find higher skin thicknesses around lap joints, as well as a minimum of two rows of fasteners at the lap joint, versus a thinner skin and only one row of fasteners through the hat section of the stringer. Any comments?
5. Last point may sound trivial but I just want to make sure: Say you're above the lap joint as described above. So you use the lap joint as the reference for obtaining your net area*Ftu and total fastener load (you take the lesser of the two). Obviously, the lap joint runs for a good length along the skin panel's length (say 100 inches). Now, your cutout is only 10 inches in length. You'd only use 10 inches of net area for your material capability, right? Likewise, when calculating the total fastener capability, you'd only multiply a single fastener's joint allowable strength (Pall) times the the number of fasteners contained above the cutout within a 10-in length segment, right?
6. Say you need to cutout part of a beam's web. Instead of using Ftu, would it be acceptable to use Fsu (much lower than Ftu) for doubler sizing since the loading in the web is primarily shear?
Hope my questions make sense. It's late. Looking forward to reading your answers.
Alex
- Status
