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  • AMERICAN BONANZA SOCIETYMID-CONTINENT AIRPORT

    WICHITA, KS 67277

    REPORT NO. ABS36-4004SA

    Revision (A)

    STRUCTURAL ANALYSIS OF FUSELAGE FRAMESAND WING CARRY-THRU STRUCTURE

    ON RAYTHEON/BEECH BONANZA TYPE AIRPLANES

    PREPARED BY:

    J.B. DWERLKOTTE ASSOCIATES, INC.429 N. ST. FRANCIS

    WICHITA, KANSAS

    DATE: September 15,2004REVISED: October 13, 2004

    /?WRITTEN BY:

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    ~y+~1I1..l2~CHECKED BY: -'11/APPROVED BY:

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 1

    1.0 Introduction:

    The purpose of this report is to analyze the reason or reasons for the continued Fuselage Frames cracking problems at the Front Wing Spar Carry-Thru Structure of the Raytheon/Beech Bonanza type airplane models as identified in Raytheon Service Bulletin SB 53-2360 Rev.(1) and the twin engine series Raytheon/Beech Travel Air/Baron models as identified in Raytheon Service Bulletin SB 53-2269.

    1.1 Description: The Single Engine Bonanza type airplanes are Type Certificated

    under (2) FAA Type Certificates, (1) A-777 for Models 35 thru G35 and including Model 35R and (2) 3A15 for Models 35-33 thru B36TC. All models in the A-777 group have Gross Weight limitations of 2850 lbs or less and all models in the 3A15 group have Gross Weights of 2900 lbs and up except the Model E33C which uses a G.W. of 2800 lbs while operating in the Acrobatic Category. All Models under the T.C.s listed above are Type Certificated in the Utility Category of CAR Part 03 and CAR Part 3 respectively. The Twin Engine Models 95 thru 58A Airplanes are all Type Certificated per T.C. 3A16 in the Normal Category of CAR Part 3 and the various models have Gross Weight limitations of 4000 lbs to 5990 lbs. The Raytheon Service Bulletins only apply to those Models listed in T.C.s 3A15 and 3A16. The above overview of the airplane models and the comparative gross weights and size of the airplanes that use the same basic wing, fuselage, and landing gear structure is provided here to give the reader a picture of similarity and small differences in loadings that occur from model to model in the review for Fatigue problems in this analysis and discussion. The Single Engine Bonanza Series of Airplanes are Type Certificated in the Utility Category and are required, by the FAA Regulations, to meet the Structural Limit Load Factor of 4.4gs, and the Twin Engine Series are Type Certificated in the Normal Category and are required to meet the Structural Limit Load Factor of not less 3.8gs, nobody flies at these load factors on purpose except in Acrobatics. Almost all flying is done at load factors of less than 2.0gs, therefore, the fatigue load factor range that is usually used and we will use in this analysis is from a minimum of 0.7gs to a maximum of 2.0gs. These load factors apply to both Flight and Landings. Any continuous flying at more than 1.5gs becomes very uncomfortable to most passengers.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 2

    1.1 Description (cont): It is important at this point for the reader to realize that the limit load factors of 4.4g's and 3.8g's respectively for the Single Engine and Twin Engine Airplanes represent the loads to which Beech Aircraft Corp. (Raytheon) was required by the FAA to demonstrate loading of the aircraft structure and hold the load on the structure for an indefinite period of time after which the load is removed and the structure returns to its original position without permanent set to any parts. No failures of parts is permitted at less than the Ultimate load factors of 6.6g's and 5.7g's respectively. Some persons such as Mr. Dick Wilson in his website "http://mysite.verizon.net" imply that the bulkheads are cracked due to high 4.4g loads. His whole analysis is based on these high "g's" which never exist except in an extreme emergency. In the real world no one flies or lands at the maximum limit load factors unless the pilot is at near emergency conditions in either flight or landing or both. Mr. Wilson's report never gets into any specific fatigue problem areas and the specific loads or range of loads that may be causing the cracks. We feel that he wants someone else to do the work for him as is indicated by his letters to Mr. Eual Conditt in the FAA ACO, Structures Section. Mr. Conditt did a very good job, in his April 12, 2004 letter, of showing the extent of the data that it takes to present and get approved an alternate means of compliance to the method presently approved for Raytheon and also it is the applicant that has the burden of supplying the supporting data. Fatigue Problems are due to stresses and deflections caused by day to day flying at lower operating loads in the 0.7g to 2.0g range that are in effect working the material in the structural parts either continuously or intermittently depending on whether the loads are caused by smooth or rough air in flight or certain landing condition loads on the main and nose gears. Any cracking in the fuselage frames is usually not a safety of flight item because the "Carry-Thru" structure, front spar and rear spar, together with the wings are complete structures without the fuselage frames being attached, however, the fuselage is not complete without attachment to the carry-thru structures for transfer of the fuselage loads into the wings. In these attachments, some of the wing differential shear loads will carry across in the frame webs. If fatigue cracks occur in the fuselage frames they should be repaired and not be allowed to remain forever. In addition, the cause of the crack's "input loads" should be positively determined so the repairs that are made will cure the problem and are not just an interim fix.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 3

    1.1 Description (cont): For the above reasons, we propose that either flight testing with strain gauges placed at predetermined critical locations and or finite element model testing as proposed in the following sections of this report should be implemented. The object of the testing is to find the "input loads" that are causing the frame webs and flanges to crack by fatigue. The Raytheon Service Bulletins allow that the first inspections for fatigue cracks is at 1500 flight hours as they determined from their customer service feed back. This shows that the loads are probably of a low magnitude but locally concentrated in the bulkhead web and skin flange area and it is highly unlikely that these are coming in from the wing thru the carry-thru attach fittings because these are all high magnitude loads and would not act like they are relieved after the bulkhead web cracks locally. As many of the service reports state, after the first cracks are found by inspection, subsequent inspections do not show a continuation of cracking at those locations. High magnitude loads in the area of cracks will always cause the cracks to grow unless reinforcements are installed to reduce the stresses and/or redirect the loads in the various parts. Understand that all loads in the airplane are continuous regardless of magnitude, therefore, if the structure in front of a load cracks, it finds another way around that crack. Therefore, the cracks in the bulkhead frame web and flange radius did not eliminate the load that caused the crack it just went around a different route in the structure. From this I think you can visualize that the loads have to be small to find a new path around the cracked location and the stress has been reduced sufficiently so that the cracked part does not continue to crack. For the above reasons, the loads causing the cracking that stopped have to be small but concentrated and the stress in the material in the localized area is in the fatigue range for the material. Figures (1A) and (1B) are scanned views from Beechcraft drawing 36-4004 Rev.(E) for information purposes to the reader only. The Current Repair Kit is divided basically into (4) Kits, (1) for each corner of the Carry-Thru Structure. Each Kit will take care of and cover the cracks in the web locally, however, all of the Kits are designed to reinforce the frames for vertical and inboard/outboard loads but they do very little for fore/aft loads which are the loads that caused the cracks in most of the photographs that we have seen. See paragraph 3.0 Analysis in this report for additional review.

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 4

    Figure (1A)

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 5

    Figure (1B)

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 6

    2.0 References: Drawings, (Beech): 36-4004 Rev.(E), Kit Information-Front Spar Carry Thru Structure Reinforcement. (For Single Engine Airplanes) 58-4008 Rev.( ), Kit Information-Front Spar carry Thru Structure Reinforcement. (For Twin Engine Airplanes) Documents: FAA AD 95-04-03, Beech Aircraft Corporation, Amendment 39-9155 (Bonanza type Single Engine Airplanes). Raytheon Service Bulletin: SB 53-2360 Rev.(1), Fuselage-Wing Forward Spar Carry-Thru Structure Inspection and/or Reinforcement. FAA AD 90-08-14, Beech Aircraft Corporation, (Small Beech Twin Engine Airplanes). Raytheon Service Bulletin: SB 53-2269 Rev.( ), Fuselage-Wing Forward Spar Carry-Thru Structure Inspection And /or Reinforcement. Raytheon Beechcraft Bonanza Series Maintenance Manuals. Photographs And Field Reports: American Bonanza Society Questionnaire Responses List (Updated 8/31/04) Photographs supplied to A.B.S. by its Members of Cracks that occurred on their Airplanes. Technical Documents: MIL-HDBK-5H Metallic Materials and Elements for Aerospace Vehicle Structures Formulas for Stress and Strain by R.J. Roark & W.C. Young Report: AFS-120-73-2 Fatigue Evaluation of Wing and Associated Structure on Small Airplanes By: FAA Engineering and Manufacturing Division Airframe Branch.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 7

    3.0 Analysis: The purpose of this analysis is to try to determine, from the

    location and type of cracks in the fuselage frames attached to the Wing Carry-Thru Structure, where the loads causing the cracks are coming from and why they are located at nearly the same quadrants on the fuselage frame in each case reported by the Field mechanics and inspectors The documented data in the American Bonanza Society Questionnaire Response List and the Photographs taken by the members and by the mechanics on location in the Field show that the primary origin of cracks is in the lower RH and LH corners of the fuselage frames near the lower RH and LH wing carry-thru fittings, which is also consistent with the Raytheon/Beech findings and the repairs by their Repair Kits shown on their drawings 36-4004 and 58-4008. The Beech Kits provide a repair that reinforces the bulkhead to skin attach flange and the bulkhead web locally where the web cracks are found near the lower fuselage to wing front spar fitting. The cracks are found alternately in the forward and aft bulkhead frames, therefore, Raytheon made Kits for each of the four corners of the Carry-Thru Structure to repair only those areas where the cracks appear. Each Kit will take care of and cover the cracks in the web locally, however, all of the Kits are designed to reinforce the frames for vertical and inboard/outbd loads but none of the Kits take care of stopping the cracking due to any fore/aft input loads. The areas of the frame flanges attached to the fuselage skin and doublers at the cutouts for the carry-thru fittings is very vulnerable to concentrated fatigue loads. A review of the photographs of the cracks found in the field and a study of the Raytheon/Beech drawing shows that most of the cracks in the bulkhead frames and the general center of the Raytheon repairs are centered on the lower front spar fittings, which passes thru an approx. 3.0 x 3.0 hole in the fuselage skin and skin doubler. In addition the fuselage skin corner J stringer is centered directly on the cutout providing a highly stiffened area in the skin locally that realizes fore/aft loads from rudder loads, unequal landing loads both main and nose gear and other side inertia loads that must be reacted around the cutouts in the fuselage skin. These loads do not have a good direct load path around the fitting cutouts and must, therefore, be bridged around the fittings by the fuselage frame flanges and fuselage skin and doublers. It has come to our attention that the 0.032 thick fuselage skin and 0.040 thick area doubler, all of 2024-T42; reference Beechcraft Bonanza Series Maintenance Manuals, may not be sufficient structure in all operating environments to carry the above named loads, that are gathered into the corner J

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 8

    3.0 Analysis (cont): stringer, around the carry-thru fittings pass-thru cutouts without slight deflections that result in passing some of the load into the bulkhead flanges attached to the skin and doublers resulting in fore/aft loads in the bulkhead web corners for which the bulkhead webs were not designed and the result is that the bulkhead web takes a finite number of cycles and then cracks. We have provided a brief analysis, to show that even small loads with the structure not fully stiffened to carry the loads around the cutouts will cause small repeated deflections that result in fatigue cracks. See Appendix (A) to this report for the analysis and see Appendix (B) for the photographs.

    The concerns outlined by Peter Harradine after his calculations in Appendix (A) of this report are concerns that need more attention by Raytheon Aircraft Services so the problems do not reappear after the Kits have been installed. The above concerns about cutouts in the fuselage skins are faced everyday in the design of pressure vessel cutouts for antennas and various other equipment on pressurized fuselages. Both upper and lower carry-thru fitting skin cutouts should be reviewed if any upgrading of the Kits is considered because cracks in the radius of the fuselage frame have also been found at the upper carry-thru fitting above the area covered by the Raytheon Kits.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 9

    4.0 Conclusion:

    The Raytheon/Beechcraft repair Kits reinforce the webs of the Fuselage Frames at the locations where the cracks have been occurring, however, we do not see that the Kits provide a permanent repair unless a set of doublers are installed on the outside of the fuselage skin to redirect the fore/aft loads around the cutouts so the fuselage frame flanges do not have to bridge the loads. The skin doublers should be installed on both sides of the fuselage regardless of whether one or four of the other Raytheon Kits are installed. It is our considered opinion that a finite element model should be made of the Fuselage, from the Rear Spar Carry-Thru forward to the Nose Landing Gear Attachment points, and the Wing, from the Landing Gear Attachment points in the RH and LH wings. Because of the similarity of the Single Engine and Twin Engine Airplanes in these areas, the application of various loading conditions to the same F.E. model will show where critical deflections occur. There should be no need for more than one F.E. model. The F.E. Model would not require finite detail of structure except in suspect areas of excess deflections. The F.E. Model should be designed with the cooperation of Raytheon to obtain good input loads, however, this is not a mandatory item.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 10

    Appendix 'A' Simplistic Deflection Analysis & Assessment of Web Stresses. by Peter Harradine, Aircraft Engineering Consultant

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    Appendix 'B' Photographs provided by Field Service and Repair Facilities.

  • AMERICAN BONANZA SOCIETY.

    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 17

    Stringer

    (RH Front Bulkhead Looking Aft)

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    (RH Front Bulkhead Looking Aft)

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 19

    Cracks Stringer

    (RH Front Bulkhead Looking Aft)

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 20

    Stringer Crack

    (RH Front Bulkhead Looking Aft)

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    REV. (A) REPORT NO. ABS36-4004SA PAGE NO. 21

    Stringer Crack

    (RH Front Bulkhead Looking Aft)