AD 90-08-14

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AD Number:
90-08-14
Document Type:
AD Final Rules
Docket Number:
89-CE-26-AD
Subject Heading:
Airworthiness Directives; Beech MODELS 95, B95, B95A, D95A, E95, 95-55, 95-A55, 95-B55, 95-B55A, 95-...Show more
Subject:
Wing Spar Structure
Status:
Current
Citation:
This information is not available.
Citation Publish Date:
Effective Date:
05/07/1990
Make:
Textron Aviation Inc.
Model:
56TC | 58 | 58A | 95 | 95-55 | 95-A55 | 95-B55 | 95-B55A | 95-B55B | 95-C55 | 95-C55A | A56TC | B95 ...Show more
Product Type:
Aircraft
Product Subtype:
Small Airplane
Affected AD:
Superseded AD:
Affected By:
Superseded By:
Service/Office:
Office of Primary Responsibility:
CFR Part Reference:
Part 39
Comments:
AIRWORTHINESS DIRECTIVES FINAL RULES: 90-08-14
CITATION:   This information is not available.

PAGE NUMBER:  

DOCKET NUMBER:   89-CE-26-AD

AMENDMENT:   39-6563

AD NUMBER:   90-08-14

SUBJECT HEADING:   Airworthiness Directives; Beech MODELS 95, B95, B95A, D95A, E95, 95-55, 95-A55, 95-B55, 95-B55A, 95-C55, 95-C55A, D55, D55A, E55, E55A, 95-B55B (T42A), 56TC, A56TC, 58, and 58A Airplanes

ACTION:  

SUMMARY:  

DATES:   Effective May 7, 1990

ADDRESSES:  

FOR FURTHER INFORMATION CONTACT:  

SUPPLEMENTARY INFORMATION:  

REGULATORY TEXT:  
90-08-14 BEECH: Amendment 39-6563. Docket No. 89-CE-26-AD.

Applicability: The following airplanes certificated in any category.

MODELS	SERIAL NUMBERS
95, B95, B95A, D95A, E95	TD-1 through TD-721
95-55, 95-A55, 95-B55 and 95-B55A	TC-1 through TC-2456, except TC-350

95-C55, 95-C55A, D55, D55A, E55 and E55A	TC-350 and TE-1 through TE-1201

95-B55B (T42A)	TF-1 through TF-70
56TC, A56TC	TG-1 through TG-94
58, 58A	TH-1 through TH-1475

Compliance: Required as indicated in the body of the AD, unless already accomplished.

To prevent cracks in the wing forward spar carry-through web structure from propagating to lengths that could compromise the integrity of the wing attachment to the fuselage, accomplish the following:

(a) Within the next 100 hours time-in-service (TIS), after the effective date of this AD, or upon the accumulation of 1,500 hours total TIS, whichever occurs later, and thereafter at the intervals specified below, inspect the wing forward spar carry-through web structure in accordance with the instructions in Beech Service Bulletin (SB) No. 2269, Revision 1, dated March 1990.

(1) If no cracks are found, repeat the inspection at 500 hour TIS intervals thereafter.

(2) For cracks in the bend radius:

(i) If the crack length is less than 2.25 inches, prior to further flight stop drill the crack in accordance with the instructions in Beech SB No. 2269, Revision 1, and reinspect for crack progression every 200 hours TIS thereafter. Only one stop drilled crack for the left side and one stop drilled crack for the right side of the web structure are permissible.

(ii) If the crack length is greater than 2.25 inches but less than 4.0 inches, prior to further flight stop drill the crack in accordance with the instructions in Beech SB No. 2269, Revision 1, and within the next 100 hours TIS, repair the web structure with the applicable Beech Part Number (P/N) 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition.

(iii) If the crack length is greater than 4.0 inches, prior to further flight repair the web structure with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition.

(3) For cracks in the web face, in the area of the huckbolt fasteners:

(i) If the crack length is less than 1.0 inch, reinspect for crack progression every 100 hours TIS thereafter. Only one crack for the left side and one crack for the right side are permissible, provided neither crack exceeds 1.0 inch in length.

NOTE 1: Do not stop drill these cracks due to the possibility of damaging the structure behind the web face.

(ii) If any crack length is greater than 1.0 inch, or a crack is connecting two fastener holes, within the next 25 hours TIS, repair the web face with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition.

(iii) If any crack passes through two fastener holes and extends beyond the holes for more than 0.5 inch, prior to further flight repair the web face with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition.

(4) If cracks are found on the same side of the airplane in both the forward and aft web face, or the bend radii, and any of the cracks are more than 1.0 inch long, prior to further flight repair the web structure with the applicable Beech P/N 58-4008 kit as specified in the above SB. After installation of the applicable Beech P/N 58-4008 kit, dye-penetrant inspect this area for cracks within the next 1,500 hours TIS from the time of installation of the applicable kit, and reinspect for cracks at 500 hours TIS intervals thereafter. If cracks are detected in these subsequent inspections, prior to further flight, contact the Wichita Aircraft Certification Office at the address below for disposition.

NOTE 2: If a fuselage skin crack is discovered around the opening for the lower forward carry-through fitting, an external doubler may be required.

(b) Airplanes may be flown in accordance with FAR 21.197 to a location where the AD may be accomplished.

(c) An alternate method of compliance or adjustment of the initial or repetitive compliance times, which provides an equivalent level of safety, may be approved by the Manager, Wichita Aircraft Certification Office, FAA, Room 100, 1801 Airport Road, Wichita, Kansas 67209.

NOTE 3: The request should be forwarded through an FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office.

All persons affected by this directive may obtain copies of the document referred to herein upon request to Beech Aircraft Corporation, Commercial Service, Department 52, P.O. Box 85, Wichita, Kansas 67201-0085; or may examine this document at the FAA, Office of the Assistant Chief Counsel, Room 1558, 601 East 12th Street, Kansas City, Missouri 64106.

This amendment (39-6563, AD 90-08-14) becomes effective on May 7, 1990.

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