AD 98-14-09
final rule
Applicability
| Type | Manufacturer | Model | Details |
|---|---|---|---|
| aircraft | The Boeing Company | 737-100 Series | Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes |
| aircraft | The Boeing Company | 737-200 Series | Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes |
| aircraft | The Boeing Company | 737-200C Series | Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes |
Unsafe Condition
Fatigue cracking of the lower flange of the forward engine mount support (FEMS) fitting, fractured or loose bolts at the upper inboard attachment, and cracked or severed lugs at the upper outboard support link attachment could result in in-flight engine separation.
AI-generated summary from the source AD text. Verify against the official source before acting.
Required Actions
Conduct repetitive detailed visual inspections and ultrasonic inspections of the FEMS fitting and its attachments to detect fatigue cracking and discrepancies. Repair or replace components as necessary, including installing newer-type FEMS fittings or serviceable older-type fittings. Report adverse findings to the FAA.
AI-generated summary from the source AD text. Verify against the official source before acting.
Compliance Time
Before further flight unless otherwise specified in the service bulletins referenced in the AD.
AI-generated summary from the source AD text. Verify against the official source before acting.
Affected Aircraft
Boeing Model 737-100, -200, and -200C Series airplanes with affected forward engine mount support (FEMS) fittings.
AI-generated summary from the source AD text. Verify against the official source before acting.
Federal Register Abstract
Forward Engine Mount Support
Applicability Source Text
Show captured applicability text from the source AD
AD Final Rules - DRS_98-14-09.html Copy URL Print/Save as PDF Document Versions Feedback Details AD Number: 98-14-09 Document Type: AD Final Rules Docket Number: 98-NM-121-AD Subject Heading: Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes Subject: Forward Engine Mount Support Status: Current Citation: (Federal Register: July 02, 1998 (Volume 63, Number 127)) Citation Publish Date: 07/02/1998 Effective Date: 07/17/1998 Make: The Boeing Company Model: 737-100 Series | 737-200 Series | 737-200C Series Product Type: Aircraft Product Subtype: Large Airplane Affected AD: Superseded AD: Affected By: Superseded By: Service/Office: Office of Primary Responsibility: CFR Part Reference: Part 39 (63 FR 36158 NO. 127 07/02/98) Comments: AIRWORTHINESS DIRECTIVES FINAL RULES: 98-14-09 CITATION: [Federal Register: July 02, 1998 (Volume 63, Number 127)] PAGE NUMBER: [Page 36158] DOCKET NUMBER: 98-NM-121-AD AMENDMENT: 39-10642 AD NUMBER: 98-14-09 SUBJECT HEADING: Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes ACTION: Final rule; request for comments SUMMARY: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Boeing Model 737-100, -200, and -200C series airplanes. This action requires repetitive inspections to detect fatigue cracking and certain discrepancies of the forward engine mount support (FEMS) fitting and its attachments, and repair, if necessary. This amendment is prompted by reports of fatigue cracks on the lower flange of the FEMS fitting, broken bolts and bolts with loose or detached nuts on the upper inboard attachment of the FEMS fitting, and cracked or severed lugs at the outboard support link attachment of the FEMS fitting. The actions specified in this AD are intended to detect and correct fatigue cracking and certain discrepancies of the FEMS fitting and its attachments, which could result in an in-flight separation of an engine. DATES: Effective July 17, 1998. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of July 17, 1998. Comments for inclusion in the Rules Docket must be received on or before August 31, 1998. ADDRESSES: Submit comments in triplicate to the Federal Aviation Administration (FAA), Transport Airplane Directorate, ANM-114, Attention: Rules Docket No. 98-NM-121-AD, 1601 Lind Avenue, SW., Renton, Washington 98055-4056. The service information referenced in this AD may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124-2207. This information may be examined at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. FOR FURTHER INFORMATION CONTACT: Gregory L. Schneider, Aerospace Engineer, Airframe Branch, ANM-120S, FAA, Transport Airplane Directorate, Seattle Aircraft Certification Office, 1601 Lind Avenue, SW., Renton, Washington 98055-4056; telephone (425) 227-2028; fax (425) 227-1181. SUPPLEMENTARY INFORMATION: The FAA has received reports of certain problems affecting the forward engine mount support (FEMS) fitting on certain Boeing Model 737 series airplanes. This support fitting is one of the primary structural elements that attach the engine to the wing. The reports indicate that three critical elements of the FEMS fitting have proved to be susceptible to fatigue damage or other problems as summarized below: Lower Flange of the FEMS Fitting: The FAA has received 17 reports of cracks of the lower flange "I" section of the FEMS fitting. Analysis indicates that the cracks were initiated by fatigue. A FEMS fitting that has a cracked lower flange may not be capable of withstanding certain limit load conditions. Upper Inboard Attachment Bolt: There have been 13 cases of the upper inboard attachment bolt fracturing in service due to fatigue, and 4 cases of the nut being broken, loose, or detached. Investigation revealed that the original production bolt installation was subject to relative motion between the bushing and the attachment bolt. As a result, the production nut (which has no secondary locking features) tended to come loose in service. A later configuration change that was intended to correct this problem consisted of installing a stronger bolt and nut, and a new bushing. This change, which has subsequently been adopted by almost the entire fleet of affected airplanes, requires the nut to be torqued to a higher value than is appropriate for the bolt and nut installation. Specifically, the torque applied to the new nut is applicable to a "non-lubricated" thread condition, whereas the nut material tends to act as a "dry" lubricant. Consequently, the higher torque applied to the new bolt and nut configuration induces an excessive pre-load on the bolt threads. This excessive pre-load, in conjunction with certain operational loads, causes an overload condition on the bolt threads, which in turn leads to premature fatigue cracking of the bolt. Additionally, results of an analysis indicate that the FEMS fitting cannot react certain limit load conditions with a fractured or detached bolt at this location. Upper Outboard Lug of the FEMS Fitting: The upper outboard lug of the FEMS fitting contains a bearing that has proved susceptible to excessive wearing. This lug is designed to secure the outboard end of the FEMS fitting to the wing. A severely worn bearing could drastically reduce the fatigue life of the lug. This condition has been observed on six airplanes to date; on three of those airplanes the lug was found to be completely fractured. Analysis has revealed that the FEMS fitting cannot react certain limit load conditions with a severed lug. Explanation of the Unsafe Condition The fatigue cracking problems that affect the three areas of the FEMS fitting are examples of "multiple element damage." The existence of any one of these conditions could result in an engine separation under certain limit load conditions. The simultaneous existence of any two conditions could result in an immediate engine loss at loads that are much lower than the design limit loads. These problems, if not corrected, could result in an in-flight separation of an engine. Explanation of Relevant Service Information The FAA has reviewed and approved the following three service bulletins: Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998, addresses fatigue cracking of the lower flange of the FEMS fitting. The service bulletin notes that the fatigue cracking problem affects only "older-type" FEMS fittings that have a lower flange thickness of 0.32 inches (nominal). Therefore, the service bulletin describes procedures for performing repetitive detailed visual inspections of the lower flange of the "older-type" FEMS fitting to detect fatigue cracking, and corrective action, if necessary. The corrective action includes replacement of the "older-type" FEMS fitting with a "newer-type" FEMS fitting, which would eliminate the need for the repetitive detailed visual inspections. These inspections are not required on "newer-type" FEMS fittings [i.e., those FEMS fittings having lower flanges that are 0.40 inches (nominal) thick], since there have been no reports of fatigue cracking of the lower flange of these parts. Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998, describes procedures for performing repetitive detailed visual inspections of the upper inboard attachment of the FEMS fitting to detect bolt deformation or fatigue damage. Additionally, the service bulletin recommends that operators perform a torque check during each inspection to ensure that the nut and bolt installation has retained its integrity. The service bulletin also describes procedures for an initial and two follow-on ultrasonic inspections of the bolt to detect fatigue cracking, and replacement of any discrepant part. The service bulletin recommends that, if the three successive ultrasonic inspections (i.e., the initial and the two follow-on inspections) reveal that the bolt is undamaged, the need for further ultrasonic inspections would be eliminated. In addition, the service bulletin describes procedures for replacement of the bolt and nut installation with a new Nickel Alloy 718 bolt and associated nut, which would eliminate the need for the repetitive detailed visual inspections and torque checks. Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998, describes procedures for repetitive detailed visual inspections of the lug of the outboard support link attachment of the FEMS fitting to detect cracked or severed lugs; and corrective action, if necessary. The service bulletin notes that some of the lug structure will not be visible during the detailed visual inspection. If a crack is detected, the corrective action is to replace the cracked FEMS fitting with a "newer-type" FEMS fitting and to install a new bearing. The service bulletin also describes procedures for an optional preventive modification, which entails removing the engine, installing a new bearing, and re-installing the existing fitting (provided that a magnetic particle inspection shows that the lug of the existing FEMS fitting is free of cracks). Explanation of the Requirements of the Rule Since an unsafe condition has been identified that is likely to exist or develop on other airplanes of the same type design, this AD is being issued to detect and correct fatigue cracking and certain discrepancies of the FEMS fitting and its attachments, which could result in an in-flight separation of an engine. This AD requires accomplishment of the actions specified in the service bulletins described previously, except as discussed below. This AD also requires that operators report any adverse (negative) inspection findings to the FAA. Differences Between the AD and the Service Bulletins Boeing Service Bulletin 737-54A1012, Revision 4, specifies that if cracking of the lower flange of the FEMS fitting is found, the cracked FEMS fitting should be replaced with a "newer-type" FEMS fitting. Such installation of a "newer-type" FEMS fitting would constitute terminating action for the repetitive detailed visual inspection requirements of this AD. However, since sufficient parts may not be available for all of the affected airplanes, this AD allows operators to install either an "older-type" FEMS fitting that is "serviceable," or a "newer-type" FEMS fitting. The installation of a "serviceable" FEMS fitting instead of a "newer-type" FEMS fitting would not terminate the repetitive detailed visual inspections required by this AD. Rather, these inspections would continue until a "newer-type" FEMS fitting is installed. For the purposes of this AD, a "serviceable" FEMS fitting is defined as an "older-type" FEMS fitting that has been shown to be free of cracks by means of a magnetic particle inspection. This AD also requires operators to perform the magnetic particle inspection in accordance with a method approved by the FAA. Although Boeing Service Bulletin 737-54-1007, Revision 1, advises operators to examine the nut of the FEMS fitting inboard attachment for looseness, it does not provide procedures for determining if the nut is too tight. This AD requires operators to examine the nut for both looseness and excessive tightness. This AD also requires that, if the nut is found to be too loose or too tight, the nut is to be re-torqued to a value of 440 to 650 pound-inches, provided that a run-on torque value of at least 18 pound-inches can be achieved. If the run-on torque value cannot be achieved, the nut is to be replaced with a new nut. This run-on torque check is to be accomplished by loosening the nut sufficiently to demonstrate that a minimum run-on torque value of 18 pound-inches can be achieved. Finally, this AD requires operators to perform this same run-on torque check on any new nut that is installed on the bolt. If a new nut should fail the 18 pound-inches minimum requirement, then this would imply that the bolt thread was defective. Therefore, if this were to occur, this AD requires the operator to replace the existing bolt installation with a stronger bolt installation in accordance with the service bulletin. Boeing Service Bulletin 737-54-1009, Revision 1, specifies that the manufacturer may be contacted for disposition of certain repair conditions (i.e., for a repair of a cracked lug). However, this AD requires that the repair of those conditions be accomplished in accordance with a method approved by the FAA. Previously Modified Airplanes Each of the three Boeing service bulletins specified in this AD contains the following statement: "If an airplane has a non-Boeing modification or repair that affects a component or system affected by this service bulletin, the operator is responsible for obtaining appropriate regulatory agency approval before incorporating this service bulletin." The FAA is aware that a certain proportion of the airplanes listed in the effectivity sections of the three service bulletins have already been modified by certain non-Boeing engine hush-kit supplemental type certificates (STC). The FAA has determined that the following hush-kit STC's are compatible with the service bulletins; therefore, operators of airplanes modified with the following STC's need not seek prior FAA approval before accomplishing the requirements of this AD. SA5730NM, issued June 26, 1992; amended October 2, 1992. ST00131SE, issued November 8, 1994; amended January 26, 1995; May 13, 1996; September 13, 1996; and February 20, 1997. ST223CH, issued July 7, 1994; amended August 11, 1994; December 19, 1994; May 30, 1995; and October 14, 1997. Interim Action This is considered to be interim action. The FAA is currently considering requiring replacement of the attachment bolt installation and the bearing with new and improved replacement parts. However, the planned compliance time for installation of new and improved parts is sufficiently long that notice and opportunity for prior public comment will be practicable. Determination of Rule's Effective Date Since a situation exists that requires the immediate adoption of this regulation, it is found that notice and opportunity for prior public comment hereon are impracticable, and that good cause exists for making this amendment effective in less than 30 days. Comments Invited Although this action is in the form of a final rule that involves requirements affecting flight safety and, thus, was not preceded by notice and an opportunity for public comment, comments are invited on this rule. Interested persons are invited to comment on this rule by submitting such written data, views, or arguments as they may desire. Communications shall identify the Rules Docket number and be submitted in triplicate to the address specified under the caption "ADDRESSES." All communications received on or before the closing date for comments will be considered, and this rule may be amended in light of the comments received. Factual information that supports the commenter's ideas and suggestions is extremely helpful in evaluating the effectiveness of the AD action and determining whether additional rulemaking action would be needed. Comments are specifically invited on the overall regulatory, economic, environmental, and energy aspects of the rule that might suggest a need to modify the rule. All comments submitted will be available, both before and after the closing date for comments, in the Rules Docket for examination by interested persons. A report that summarizes each FAA-public contact concerned with the substance of this AD will be filed in the Rules Docket. Commenters wishing the FAA to acknowledge receipt of their comments submitted in response to this rule must submit a self-addressed, stamped postcard on which the following statement is made: "Comments to Docket Number 98-NM-121-AD." The postcard will be date stamped and returned to the commenter. Regulatory Impact The regulations adopted herein will not have substantial direct effects on the States, on the relationship between the national government and the States, or on the distribution of power and responsibilities among the various levels of government. Therefore, in accordance with Executive Order 12612, it is determined that this final rule does not have sufficient federalism implications to warrant the preparation of a Federalism Assessment. The FAA has determined that this regulation is an emergency regulation that must be issued immediately to correct an unsafe condition in aircraft, and that it is not a "significant regulatory action" under Executive Order 12866. It has been determined further that this action involves an emergency regulation under DOT Regulatory Policies and Procedures (44 FR 11034, February 26, 1979). If it is determined that this emergency regulation otherwise would be significant under DOT Regulatory Policies and Procedures, a final regulatory evaluation will be prepared and placed in the Rules Docket. A copy of it, if filed, may be obtained from the Rules Docket at the location provided under the caption "ADDRESSES." List of Subjects in 14 CFR Part 39 Air transportation, Aircraft, Aviation safety, Incorporation by reference, Safety. Adoption of the Amendment Accordingly, pursuant to the authority delegated to me by the Administrator, the Federal Aviation Administration amends part 39 of the Federal Aviation Regulations (14 CFR part 39) as follows: PART 39 - AIRWORTHINESS DIRECTIVES 1. The authority citation for part 39 continues to read as follows: Authority: 49 U.S.C. 106(g), 40113, 44701. § 39.13 [Amended] 2. Section 39.13 is amended by adding the following new airworthiness directive: REGULATORY TEXT: 98-14-09 BOEING: Amendment 39-10642. Docket 98-NM-121-AD. Applicability: Model 737-100, -200, -200C series airplanes, manufacturer's line positions 001 through 1585 inclusive; certificated in any category. NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (e) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. NOTE 2: The performance of the requirements of this AD is not affected by modifications in accordance with the following supplemental type certificates (STC's). SA5730NM, issued June 26, 1992; amended October 2, 1992. ST00131SE, issued November 8, 1994; amended January 26, 1995; May 13, 1996; September 13, 1996; and February 20, 1997. ST223CH, issued July 7, 1994; amended August 11, 1994; December 19, 1994; May 30, 1995; and October 14, 1997. Compliance: Required as indicated, unless accomplished previously. To detect and correct fatigue cracking and certain discrepancies of the forward engine mount support (FEMS) fitting and its attachments, which could result in an in-flight separation of an engine, accomplish the following: (a) For airplanes on which a "newer-type" FEMS fitting having part number (P/N) 65-46850-9/-10 or 65-46850-13/-14 has not been installed: Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect fatigue cracking of the lower flange of the FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998. (1) If no fatigue cracking of the lower flange of the FEMS fitting is found, or if a "serviceable" FEMS fitting is installed in lieu of a "newer-type" FEMS fitting, repeat the inspection thereafter at intervals not to exceed 700 flight cycles in accordance with the service bulletin. NOTE 3: For the purposes of this AD, a "serviceable" FEMS fitting is defined as an "older-type" FEMS fitting that is free of cracking, as shown by a magnetic particle inspection performed in accordance with a method approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. (2) If any cracking of the lower flange of the FEMS fitting is found, prior to further flight, replace the FEMS fitting with a "serviceable" or a "newer-type" FEMS fitting in accordance with the service bulletin. Replacement of this part with a "newer-type" FEMS fitting constitutes terminating action for the repetitive inspection requirements of paragraph (a)(1) of this AD. (b) Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect deformation or fatigue damage of the bolt at the upper inboard attachment of the FEMS fitting; perform a torque check to detect any bolt that is under- or over-torqued; and perform an ultrasonic inspection to detect any cracking of the bolt; in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998. (1) If no bolt deformation or fatigue damage, under- or over-torqued nut, or fatigue cracking is found: Thereafter, repeat the detailed visual inspection and torque check required by paragraph (b) of this AD at intervals not to exceed 700 flight cycles. Additionally, repeat the ultrasonic inspection two more times at intervals not to exceed 700 flight cycles, but no earlier than 600 flight cycles. (2) If any deformation, fatigue damage, or fatigue cracking of the inboard attachment bolt is found during any inspection required by this paragraph: Prior to further flight, replace the inboard attachment bolt and nut with a new Nickel Alloy 718 bolt and associated nut in accordance with the service bulletin. Replacement of the inboard attachment bolt and nut in accordance with the service bulletin constitutes terminating action for the repetitive inspection requirements of paragraphs (b)(1), (b)(2), and (b)(3) of this AD. (3) If the torque check shows that a nut is torqued to any value outside the limits of 440 to 650 pound-inches, prior to further flight, accomplish paragraphs (b)(3)(i) and (b)(3)(ii) of this AD. (i) Loosen the affected nut enough to demonstrate that a minimum run-on torque value of 18 pound-inches can be achieved. If this value cannot be achieved, install a new nut in accordance with the service bulletin, and repeat the run-on torque check prior to tightening the nut to 440-650 inch pounds. If a run-on torque value of 18 pound-inches still cannot be achieved, prior to further flight, replace the inboard attachment bolt and nut with a new Nickel Alloy 718 bolt and associated nut in accordance with the service bulletin. (ii) Tighten the affected nut to 440-650 pound-inches in accordance with the service bulletin. (c) Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect any cracked or severed lug of the outboard support link attachment of the FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. (1) If no cracked or severed lug is detected: Repeat the detailed visual inspection required by paragraph (c) thereafter at intervals not to exceed 700 flight cycles, or perform the optional terminating modification, in accordance with Part II of the Accomplishment Instructions of the service bulletin. Where the service bulletin specifies that the manufacturer may be contacted for disposition of certain repair conditions, repair in accordance with a method approved by the Manager, Seattle ACO. Accomplishment of this modification constitutes terminating action for the repetitive inspection requirements of paragraph (c) of this AD. (2) If any cracked or severed lug is found, prior to further flight, accomplish the requirements of paragraphs (c)(2)(i) and (c)(2)(ii) of this AD. (i) Replace the FEMS fitting with a "serviceable" or a "newer-type" FEMS fitting in accordance with Accomplishment Instructions of Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998. Replacement of the FEMS fitting with a "newer-type" FEMS fitting in accordance with the service bulletin constitutes terminating action for the repetitive inspection requirements of paragraph (a) of this AD. (ii) Install a new bearing, which is inserted into the lug of the replacement FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. Replacement of the existing bearing with an improved bearing constitutes terminating action for the repetitive inspection requirements of the lug that are specified in paragraph (c) of this AD. (d) Within 20 days after accomplishing the initial inspections required by paragraphs (a), (b), and (c) of this AD, or within 20 days after the effective date of this AD, whichever occurs later, submit a report of the inspection results (adverse findings only) to the Manager, Seattle ACO, FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington 98055-4056; fax (425) 227-1181. Required information for each report must include the following: A description of the adverse finding, airplane serial number and total flight cycles and flight hours accumulated, number of flight cycles and flight hours accumulated since the last engine change, and the number of flight cycles and flight hours accumulated since the last inspection of the affected part. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (44 U.S.C. 3501 et seq.) and have been assigned OMB Control Number 2120-0056. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Seattle ACO. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Seattle ACO. NOTE 4: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Seattle ACO. (f) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. (g) Except as provided in paragraph (c)(1) of this AD, the actions shall be done in accordance with Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998; Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998; and Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124-2207. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (h) This amendment becomes effective on July 17, 1998. FOOTER:
Document Text
Show stored source text (verify against official source)
AD Final Rules - DRS_98-14-09.html Copy URL Print/Save as PDF Document Versions Feedback Details AD Number: 98-14-09 Document Type: AD Final Rules Docket Number: 98-NM-121-AD Subject Heading: Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes Subject: Forward Engine Mount Support Status: Current Citation: (Federal Register: July 02, 1998 (Volume 63, Number 127)) Citation Publish Date: 07/02/1998 Effective Date: 07/17/1998 Make: The Boeing Company Model: 737-100 Series | 737-200 Series | 737-200C Series Product Type: Aircraft Product Subtype: Large Airplane Affected AD: Superseded AD: Affected By: Superseded By: Service/Office: Office of Primary Responsibility: CFR Part Reference: Part 39 (63 FR 36158 NO. 127 07/02/98) Comments: AIRWORTHINESS DIRECTIVES FINAL RULES: 98-14-09 CITATION: [Federal Register: July 02, 1998 (Volume 63, Number 127)] PAGE NUMBER: [Page 36158] DOCKET NUMBER: 98-NM-121-AD AMENDMENT: 39-10642 AD NUMBER: 98-14-09 SUBJECT HEADING: Airworthiness Directives; Boeing Model 737-100, -200, -200C Series Airplanes ACTION: Final rule; request for comments SUMMARY: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Boeing Model 737-100, -200, and -200C series airplanes. This action requires repetitive inspections to detect fatigue cracking and certain discrepancies of the forward engine mount support (FEMS) fitting and its attachments, and repair, if necessary. This amendment is prompted by reports of fatigue cracks on the lower flange of the FEMS fitting, broken bolts and bolts with loose or detached nuts on the upper inboard attachment of the FEMS fitting, and cracked or severed lugs at the outboard support link attachment of the FEMS fitting. The actions specified in this AD are intended to detect and correct fatigue cracking and certain discrepancies of the FEMS fitting and its attachments, which could result in an in-flight separation of an engine. DATES: Effective July 17, 1998. The incorporation by reference of certain publications listed in the regulations is approved by the Director of the Federal Register as of July 17, 1998. Comments for inclusion in the Rules Docket must be received on or before August 31, 1998. ADDRESSES: Submit comments in triplicate to the Federal Aviation Administration (FAA), Transport Airplane Directorate, ANM-114, Attention: Rules Docket No. 98-NM-121-AD, 1601 Lind Avenue, SW., Renton, Washington 98055-4056. The service information referenced in this AD may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124-2207. This information may be examined at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. FOR FURTHER INFORMATION CONTACT: Gregory L. Schneider, Aerospace Engineer, Airframe Branch, ANM-120S, FAA, Transport Airplane Directorate, Seattle Aircraft Certification Office, 1601 Lind Avenue, SW., Renton, Washington 98055-4056; telephone (425) 227-2028; fax (425) 227-1181. SUPPLEMENTARY INFORMATION: The FAA has received reports of certain problems affecting the forward engine mount support (FEMS) fitting on certain Boeing Model 737 series airplanes. This support fitting is one of the primary structural elements that attach the engine to the wing. The reports indicate that three critical elements of the FEMS fitting have proved to be susceptible to fatigue damage or other problems as summarized below: Lower Flange of the FEMS Fitting: The FAA has received 17 reports of cracks of the lower flange "I" section of the FEMS fitting. Analysis indicates that the cracks were initiated by fatigue. A FEMS fitting that has a cracked lower flange may not be capable of withstanding certain limit load conditions. Upper Inboard Attachment Bolt: There have been 13 cases of the upper inboard attachment bolt fracturing in service due to fatigue, and 4 cases of the nut being broken, loose, or detached. Investigation revealed that the original production bolt installation was subject to relative motion between the bushing and the attachment bolt. As a result, the production nut (which has no secondary locking features) tended to come loose in service. A later configuration change that was intended to correct this problem consisted of installing a stronger bolt and nut, and a new bushing. This change, which has subsequently been adopted by almost the entire fleet of affected airplanes, requires the nut to be torqued to a higher value than is appropriate for the bolt and nut installation. Specifically, the torque applied to the new nut is applicable to a "non-lubricated" thread condition, whereas the nut material tends to act as a "dry" lubricant. Consequently, the higher torque applied to the new bolt and nut configuration induces an excessive pre-load on the bolt threads. This excessive pre-load, in conjunction with certain operational loads, causes an overload condition on the bolt threads, which in turn leads to premature fatigue cracking of the bolt. Additionally, results of an analysis indicate that the FEMS fitting cannot react certain limit load conditions with a fractured or detached bolt at this location. Upper Outboard Lug of the FEMS Fitting: The upper outboard lug of the FEMS fitting contains a bearing that has proved susceptible to excessive wearing. This lug is designed to secure the outboard end of the FEMS fitting to the wing. A severely worn bearing could drastically reduce the fatigue life of the lug. This condition has been observed on six airplanes to date; on three of those airplanes the lug was found to be completely fractured. Analysis has revealed that the FEMS fitting cannot react certain limit load conditions with a severed lug. Explanation of the Unsafe Condition The fatigue cracking problems that affect the three areas of the FEMS fitting are examples of "multiple element damage." The existence of any one of these conditions could result in an engine separation under certain limit load conditions. The simultaneous existence of any two conditions could result in an immediate engine loss at loads that are much lower than the design limit loads. These problems, if not corrected, could result in an in-flight separation of an engine. Explanation of Relevant Service Information The FAA has reviewed and approved the following three service bulletins: Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998, addresses fatigue cracking of the lower flange of the FEMS fitting. The service bulletin notes that the fatigue cracking problem affects only "older-type" FEMS fittings that have a lower flange thickness of 0.32 inches (nominal). Therefore, the service bulletin describes procedures for performing repetitive detailed visual inspections of the lower flange of the "older-type" FEMS fitting to detect fatigue cracking, and corrective action, if necessary. The corrective action includes replacement of the "older-type" FEMS fitting with a "newer-type" FEMS fitting, which would eliminate the need for the repetitive detailed visual inspections. These inspections are not required on "newer-type" FEMS fittings [i.e., those FEMS fittings having lower flanges that are 0.40 inches (nominal) thick], since there have been no reports of fatigue cracking of the lower flange of these parts. Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998, describes procedures for performing repetitive detailed visual inspections of the upper inboard attachment of the FEMS fitting to detect bolt deformation or fatigue damage. Additionally, the service bulletin recommends that operators perform a torque check during each inspection to ensure that the nut and bolt installation has retained its integrity. The service bulletin also describes procedures for an initial and two follow-on ultrasonic inspections of the bolt to detect fatigue cracking, and replacement of any discrepant part. The service bulletin recommends that, if the three successive ultrasonic inspections (i.e., the initial and the two follow-on inspections) reveal that the bolt is undamaged, the need for further ultrasonic inspections would be eliminated. In addition, the service bulletin describes procedures for replacement of the bolt and nut installation with a new Nickel Alloy 718 bolt and associated nut, which would eliminate the need for the repetitive detailed visual inspections and torque checks. Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998, describes procedures for repetitive detailed visual inspections of the lug of the outboard support link attachment of the FEMS fitting to detect cracked or severed lugs; and corrective action, if necessary. The service bulletin notes that some of the lug structure will not be visible during the detailed visual inspection. If a crack is detected, the corrective action is to replace the cracked FEMS fitting with a "newer-type" FEMS fitting and to install a new bearing. The service bulletin also describes procedures for an optional preventive modification, which entails removing the engine, installing a new bearing, and re-installing the existing fitting (provided that a magnetic particle inspection shows that the lug of the existing FEMS fitting is free of cracks). Explanation of the Requirements of the Rule Since an unsafe condition has been identified that is likely to exist or develop on other airplanes of the same type design, this AD is being issued to detect and correct fatigue cracking and certain discrepancies of the FEMS fitting and its attachments, which could result in an in-flight separation of an engine. This AD requires accomplishment of the actions specified in the service bulletins described previously, except as discussed below. This AD also requires that operators report any adverse (negative) inspection findings to the FAA. Differences Between the AD and the Service Bulletins Boeing Service Bulletin 737-54A1012, Revision 4, specifies that if cracking of the lower flange of the FEMS fitting is found, the cracked FEMS fitting should be replaced with a "newer-type" FEMS fitting. Such installation of a "newer-type" FEMS fitting would constitute terminating action for the repetitive detailed visual inspection requirements of this AD. However, since sufficient parts may not be available for all of the affected airplanes, this AD allows operators to install either an "older-type" FEMS fitting that is "serviceable," or a "newer-type" FEMS fitting. The installation of a "serviceable" FEMS fitting instead of a "newer-type" FEMS fitting would not terminate the repetitive detailed visual inspections required by this AD. Rather, these inspections would continue until a "newer-type" FEMS fitting is installed. For the purposes of this AD, a "serviceable" FEMS fitting is defined as an "older-type" FEMS fitting that has been shown to be free of cracks by means of a magnetic particle inspection. This AD also requires operators to perform the magnetic particle inspection in accordance with a method approved by the FAA. Although Boeing Service Bulletin 737-54-1007, Revision 1, advises operators to examine the nut of the FEMS fitting inboard attachment for looseness, it does not provide procedures for determining if the nut is too tight. This AD requires operators to examine the nut for both looseness and excessive tightness. This AD also requires that, if the nut is found to be too loose or too tight, the nut is to be re-torqued to a value of 440 to 650 pound-inches, provided that a run-on torque value of at least 18 pound-inches can be achieved. If the run-on torque value cannot be achieved, the nut is to be replaced with a new nut. This run-on torque check is to be accomplished by loosening the nut sufficiently to demonstrate that a minimum run-on torque value of 18 pound-inches can be achieved. Finally, this AD requires operators to perform this same run-on torque check on any new nut that is installed on the bolt. If a new nut should fail the 18 pound-inches minimum requirement, then this would imply that the bolt thread was defective. Therefore, if this were to occur, this AD requires the operator to replace the existing bolt installation with a stronger bolt installation in accordance with the service bulletin. Boeing Service Bulletin 737-54-1009, Revision 1, specifies that the manufacturer may be contacted for disposition of certain repair conditions (i.e., for a repair of a cracked lug). However, this AD requires that the repair of those conditions be accomplished in accordance with a method approved by the FAA. Previously Modified Airplanes Each of the three Boeing service bulletins specified in this AD contains the following statement: "If an airplane has a non-Boeing modification or repair that affects a component or system affected by this service bulletin, the operator is responsible for obtaining appropriate regulatory agency approval before incorporating this service bulletin." The FAA is aware that a certain proportion of the airplanes listed in the effectivity sections of the three service bulletins have already been modified by certain non-Boeing engine hush-kit supplemental type certificates (STC). The FAA has determined that the following hush-kit STC's are compatible with the service bulletins; therefore, operators of airplanes modified with the following STC's need not seek prior FAA approval before accomplishing the requirements of this AD. SA5730NM, issued June 26, 1992; amended October 2, 1992. ST00131SE, issued November 8, 1994; amended January 26, 1995; May 13, 1996; September 13, 1996; and February 20, 1997. ST223CH, issued July 7, 1994; amended August 11, 1994; December 19, 1994; May 30, 1995; and October 14, 1997. Interim Action This is considered to be interim action. The FAA is currently considering requiring replacement of the attachment bolt installation and the bearing with new and improved replacement parts. However, the planned compliance time for installation of new and improved parts is sufficiently long that notice and opportunity for prior public comment will be practicable. Determination of Rule's Effective Date Since a situation exists that requires the immediate adoption of this regulation, it is found that notice and opportunity for prior public comment hereon are impracticable, and that good cause exists for making this amendment effective in less than 30 days. Comments Invited Although this action is in the form of a final rule that involves requirements affecting flight safety and, thus, was not preceded by notice and an opportunity for public comment, comments are invited on this rule. Interested persons are invited to comment on this rule by submitting such written data, views, or arguments as they may desire. Communications shall identify the Rules Docket number and be submitted in triplicate to the address specified under the caption "ADDRESSES." All communications received on or before the closing date for comments will be considered, and this rule may be amended in light of the comments received. Factual information that supports the commenter's ideas and suggestions is extremely helpful in evaluating the effectiveness of the AD action and determining whether additional rulemaking action would be needed. Comments are specifically invited on the overall regulatory, economic, environmental, and energy aspects of the rule that might suggest a need to modify the rule. All comments submitted will be available, both before and after the closing date for comments, in the Rules Docket for examination by interested persons. A report that summarizes each FAA-public contact concerned with the substance of this AD will be filed in the Rules Docket. Commenters wishing the FAA to acknowledge receipt of their comments submitted in response to this rule must submit a self-addressed, stamped postcard on which the following statement is made: "Comments to Docket Number 98-NM-121-AD." The postcard will be date stamped and returned to the commenter. Regulatory Impact The regulations adopted herein will not have substantial direct effects on the States, on the relationship between the national government and the States, or on the distribution of power and responsibilities among the various levels of government. Therefore, in accordance with Executive Order 12612, it is determined that this final rule does not have sufficient federalism implications to warrant the preparation of a Federalism Assessment. The FAA has determined that this regulation is an emergency regulation that must be issued immediately to correct an unsafe condition in aircraft, and that it is not a "significant regulatory action" under Executive Order 12866. It has been determined further that this action involves an emergency regulation under DOT Regulatory Policies and Procedures (44 FR 11034, February 26, 1979). If it is determined that this emergency regulation otherwise would be significant under DOT Regulatory Policies and Procedures, a final regulatory evaluation will be prepared and placed in the Rules Docket. A copy of it, if filed, may be obtained from the Rules Docket at the location provided under the caption "ADDRESSES." List of Subjects in 14 CFR Part 39 Air transportation, Aircraft, Aviation safety, Incorporation by reference, Safety. Adoption of the Amendment Accordingly, pursuant to the authority delegated to me by the Administrator, the Federal Aviation Administration amends part 39 of the Federal Aviation Regulations (14 CFR part 39) as follows: PART 39 - AIRWORTHINESS DIRECTIVES 1. The authority citation for part 39 continues to read as follows: Authority: 49 U.S.C. 106(g), 40113, 44701. § 39.13 [Amended] 2. Section 39.13 is amended by adding the following new airworthiness directive: REGULATORY TEXT: 98-14-09 BOEING: Amendment 39-10642. Docket 98-NM-121-AD. Applicability: Model 737-100, -200, -200C series airplanes, manufacturer's line positions 001 through 1585 inclusive; certificated in any category. NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (e) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. NOTE 2: The performance of the requirements of this AD is not affected by modifications in accordance with the following supplemental type certificates (STC's). SA5730NM, issued June 26, 1992; amended October 2, 1992. ST00131SE, issued November 8, 1994; amended January 26, 1995; May 13, 1996; September 13, 1996; and February 20, 1997. ST223CH, issued July 7, 1994; amended August 11, 1994; December 19, 1994; May 30, 1995; and October 14, 1997. Compliance: Required as indicated, unless accomplished previously. To detect and correct fatigue cracking and certain discrepancies of the forward engine mount support (FEMS) fitting and its attachments, which could result in an in-flight separation of an engine, accomplish the following: (a) For airplanes on which a "newer-type" FEMS fitting having part number (P/N) 65-46850-9/-10 or 65-46850-13/-14 has not been installed: Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect fatigue cracking of the lower flange of the FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998. (1) If no fatigue cracking of the lower flange of the FEMS fitting is found, or if a "serviceable" FEMS fitting is installed in lieu of a "newer-type" FEMS fitting, repeat the inspection thereafter at intervals not to exceed 700 flight cycles in accordance with the service bulletin. NOTE 3: For the purposes of this AD, a "serviceable" FEMS fitting is defined as an "older-type" FEMS fitting that is free of cracking, as shown by a magnetic particle inspection performed in accordance with a method approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. (2) If any cracking of the lower flange of the FEMS fitting is found, prior to further flight, replace the FEMS fitting with a "serviceable" or a "newer-type" FEMS fitting in accordance with the service bulletin. Replacement of this part with a "newer-type" FEMS fitting constitutes terminating action for the repetitive inspection requirements of paragraph (a)(1) of this AD. (b) Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect deformation or fatigue damage of the bolt at the upper inboard attachment of the FEMS fitting; perform a torque check to detect any bolt that is under- or over-torqued; and perform an ultrasonic inspection to detect any cracking of the bolt; in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998. (1) If no bolt deformation or fatigue damage, under- or over-torqued nut, or fatigue cracking is found: Thereafter, repeat the detailed visual inspection and torque check required by paragraph (b) of this AD at intervals not to exceed 700 flight cycles. Additionally, repeat the ultrasonic inspection two more times at intervals not to exceed 700 flight cycles, but no earlier than 600 flight cycles. (2) If any deformation, fatigue damage, or fatigue cracking of the inboard attachment bolt is found during any inspection required by this paragraph: Prior to further flight, replace the inboard attachment bolt and nut with a new Nickel Alloy 718 bolt and associated nut in accordance with the service bulletin. Replacement of the inboard attachment bolt and nut in accordance with the service bulletin constitutes terminating action for the repetitive inspection requirements of paragraphs (b)(1), (b)(2), and (b)(3) of this AD. (3) If the torque check shows that a nut is torqued to any value outside the limits of 440 to 650 pound-inches, prior to further flight, accomplish paragraphs (b)(3)(i) and (b)(3)(ii) of this AD. (i) Loosen the affected nut enough to demonstrate that a minimum run-on torque value of 18 pound-inches can be achieved. If this value cannot be achieved, install a new nut in accordance with the service bulletin, and repeat the run-on torque check prior to tightening the nut to 440-650 inch pounds. If a run-on torque value of 18 pound-inches still cannot be achieved, prior to further flight, replace the inboard attachment bolt and nut with a new Nickel Alloy 718 bolt and associated nut in accordance with the service bulletin. (ii) Tighten the affected nut to 440-650 pound-inches in accordance with the service bulletin. (c) Within 90 days or 700 flight cycles after the effective date of this AD, whichever occurs later, perform a detailed visual inspection to detect any cracked or severed lug of the outboard support link attachment of the FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. (1) If no cracked or severed lug is detected: Repeat the detailed visual inspection required by paragraph (c) thereafter at intervals not to exceed 700 flight cycles, or perform the optional terminating modification, in accordance with Part II of the Accomplishment Instructions of the service bulletin. Where the service bulletin specifies that the manufacturer may be contacted for disposition of certain repair conditions, repair in accordance with a method approved by the Manager, Seattle ACO. Accomplishment of this modification constitutes terminating action for the repetitive inspection requirements of paragraph (c) of this AD. (2) If any cracked or severed lug is found, prior to further flight, accomplish the requirements of paragraphs (c)(2)(i) and (c)(2)(ii) of this AD. (i) Replace the FEMS fitting with a "serviceable" or a "newer-type" FEMS fitting in accordance with Accomplishment Instructions of Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998. Replacement of the FEMS fitting with a "newer-type" FEMS fitting in accordance with the service bulletin constitutes terminating action for the repetitive inspection requirements of paragraph (a) of this AD. (ii) Install a new bearing, which is inserted into the lug of the replacement FEMS fitting, in accordance with the Accomplishment Instructions of Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. Replacement of the existing bearing with an improved bearing constitutes terminating action for the repetitive inspection requirements of the lug that are specified in paragraph (c) of this AD. (d) Within 20 days after accomplishing the initial inspections required by paragraphs (a), (b), and (c) of this AD, or within 20 days after the effective date of this AD, whichever occurs later, submit a report of the inspection results (adverse findings only) to the Manager, Seattle ACO, FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington 98055-4056; fax (425) 227-1181. Required information for each report must include the following: A description of the adverse finding, airplane serial number and total flight cycles and flight hours accumulated, number of flight cycles and flight hours accumulated since the last engine change, and the number of flight cycles and flight hours accumulated since the last inspection of the affected part. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provisions of the Paperwork Reduction Act of 1980 (44 U.S.C. 3501 et seq.) and have been assigned OMB Control Number 2120-0056. (e) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Seattle ACO. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Seattle ACO. NOTE 4: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Seattle ACO. (f) Special flight permits may be issued in accordance with sections 21.197 and 21.199 of the Federal Aviation Regulations (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. (g) Except as provided in paragraph (c)(1) of this AD, the actions shall be done in accordance with Boeing Service Bulletin 737-54A1012, Revision 4, dated March 26, 1998; Boeing Service Bulletin 737-54-1007, Revision 1, dated March 26, 1998; and Boeing Service Bulletin 737-54-1009, Revision 1, dated March 26, 1998. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. Copies may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124-2207. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (h) This amendment becomes effective on July 17, 1998. FOOTER:
Source: Official FAA Source ↗
Retrieved: Apr 8, 2026
Rights: U.S. Government Public Domain
This site is not affiliated with or endorsed by the FAA. Always verify with official sources.
Related Airworthiness Directives
2025-02-02 Recurring
final ruleAirworthiness Directives; The Boeing Company Airplanes
Unsafe Condition: cracks in the frame inner chord and web at station (STA) 727 between S-11L and S-13L, which could result in the inability of a principal structural element to sustain limit load, and adversely affe...
2023-13-12 Recurring
final ruleAirworthiness Directives; The Boeing Company Airplanes
Unsafe Condition: premature aging of certain passenger chemical oxygen generators, which could fail to deliver oxygen during an emergency, possibly resulting in injury to the airplane occupants.
2023-12-13
final ruleAirworthiness Directives; The Boeing Company Airplanes
Unsafe Condition: Radio altimeters cannot be relied upon to perform their intended function if they experience interference from wireless broadband operations in the 5G C-Band. This interference can cause other airp...
2023-10-02
final ruleAirworthiness Directives; Transport and Commuter Category Airplanes
Unsafe Condition: Radio altimeter anomalies that are undetected by the automation or pilot, particularly close to the ground (e.g., landing flare), could lead to loss of continued safe flight and landing. Additional...