section_id,title_number,title_name,chapter,subchapter,part_number,part_name,subpart,subpart_name,section_number,section_heading,agency,authority,source_citation,amendment_citations,full_text 10:10:1.0.1.1.18.0.81.1,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.1 Purpose.,NRC,,,"[62 FR 17687, Apr. 11, 1997]","The regulations in this part establish procedures for granting, reinstating, extending, transferring, and terminating access authorizations of licensee personnel, licensee contractors or agents, and other persons (e.g., individuals involved in adjudicatory procedures as set forth in 10 CFR part 2, subpart I) who may require access to classified information." 10:10:1.0.1.1.18.0.81.2,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.3 Scope.,NRC,,,"[70 FR 32227, June 2, 2005]","The regulations in this part apply to licensees, certificate holders, and others who may require access to classified information related to a license, certificate, an application for a license or certificate, or other activities as the Commission may determine." 10:10:1.0.1.1.18.0.81.3,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.5 Definitions.,NRC,,,"[45 FR 14481, Mar. 5, 1980, as amended at 46 FR 58283, Dec. 1, 1981; 47 FR 38683, Sept. 2, 1982; 48 FR 24320, June 1, 1983; 50 FR 36984, Sept. 11, 1985; 55 FR 11574, Mar. 29, 1990; 62 FR 17687, Apr. 11, 1997; 64 FR 15647, Apr. 1, 1999; 70 FR 32227, June 2, 2005; 75 FR 73941, Nov. 30, 2010; 86 FR 43401, Aug. 9, 2021; 87 FR 45241, July 28, 2022]","Access authorization means an administrative determination that an individual (including a consultant) who is employed by or an applicant for employment with the NRC, NRC contractors, agents, licensees and certificate holders, or other person designated by the Executive Director for Operations, is eligible for a security clearance for access to classified information. Act means the Atomic Energy Act of 1954 (68 Stat. 919), as amended. Certificate holder means a facility operating under the provisions of parts 71 or 76 of this chapter. Classified information means either classified National Security Information, Restricted Data, or Formerly Restricted Data or any one of them. It is the generic term for information requiring protection in the interest of National Security whether classified under an Executive Order or the Atomic Energy Act. Classified National Security Information means information that has been determined under E.O. 13526, as amended, or any predecessor or successor order to require protection against unauthorized disclosure and that is so designated. Cognizant Security Agency (CSA) means agencies of the Executive Branch that have been authorized by E.O. 12829 to establish an industrial security program for the purpose of safeguarding classified information under the jurisdiction of those agencies when disclosed or released to U.S. industry. These agencies are the Department of Defense, the Department of Energy, the Central Intelligence Agency, and the Nuclear Regulatory Commission. A facility has a single CSA which exercises primary authority for the protection of classified information at the facility. The CSA for the facility provides security representation for other government agencies with security interests at the facility. The Secretary of Defense has been designated as Executive Agent for the National Industrial Security Program. Commission means the Nuclear Regulatory Commission or its duly authorized representatives. “L” access authorization means an access authorization granted by the Commission that is normally based on a Tier 3 (T3) investigation conducted by the Defense Counterintelligence and Security Agency (DCSA). License means a license issued pursuant to 10 CFR parts 50, 52, 60, 63, 70, or 72. Matter means documents or material. National Security Information means information that has been determined pursuant to Executive Order 12958, as amended, or any predecessor order to require protection against unauthorized disclosure and that is so designated. Need-to-know means a determination made by an authorized holder of classified information that a prospective recipient requires access to a specific classified information to perform or assist in a lawful and authorized governmental function under the cognizance of the Commission. Person means (1) any individual, corporation, partnership, firm, association, trust, estate, public or private institution, group, government agency other than the Commission or the Department of Energy (DOE), except that the DOE shall be considered a person to the extent that its facilities are subject to the licensing and related regulatory authority of the Commission pursuant to section 202 of the Energy Reorganization Act of 1974 and sections 104, 105 and 202 of the Uranium Mill Tailings Radiation Control Act of 1978, any State or any political subdivision of, or any political entity within a State, any foreign government or nation or any political subdivision of any such government or nation, or other entity; and (2) any legal successor, representative, agent, or agency of the foregoing. “Q” access authorization means an access authorization granted by the Commission normally based on a Tier 5 (T5) investigation conducted by the Defense Counterintelligence and Security Agency, the Federal Bureau of Investigation, or other U.S. Government agency that conducts personnel security investigations. Restricted Data means all data concerning design, manufacture or utilization of atomic weapons, the production of special nuclear material, or the use of special nuclear material in the production of energy, but shall not include data declassified or removed from the Restricted Data category pursuant to section 142 of the Act. Visit authorization letters (VAL) means a letter, generated by a licensee, certificate holder or other organization under the requirements of 10 CFR parts 25 and/or 95, verifying the need-to-know and access authorization of an individual from that organization who needs to visit another authorized facility for the purpose of exchanging or acquiring classified information related to the license." 10:10:1.0.1.1.18.0.81.4,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.7 Interpretations.,NRC,,,"[45 FR 14481, Mar. 5, 1980, as amended at 90 FR 55628, Dec. 3, 2025]","Except as specifically authorized by the Commission in writing, no interpretation of the meaning of the regulations in this part by any officer or employee of the Commission other than a written interpretation by the General Counsel will be recognized to be binding upon the Commission. This section shall cease to have effect on January 8, 2027, unless the NRC determines that the cessation deadline should be extended to a date not more than 5 years in the future after offering the public an opportunity to provide input on the costs and benefits of this section and considering that input. The NRC will publish a document in the Federal Register announcing its determination and revising or removing this section accordingly." 10:10:1.0.1.1.18.0.81.5,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.8 Information collection requirements: OMB approval.,NRC,,,"[49 FR 19624, May 9, 1984, as amended at 57 FR 3720, Jan. 31, 1992; 62 FR 17687, Apr. 11, 1997; 62 FR 52185, Oct. 6, 1997; 87 FR 45241, July 28, 2022]","(a) The Nuclear Regulatory Commission has submitted the information collection requirements contained in this part to the Office of Management and Budget (OMB) for approval as required by the Paperwork Reduction Act (44 U.S.C. 3501 et seq. ). The NRC may not conduct or sponsor and a person is not required to respond to, a collection of information unless it displays a currently valid OMB control number. OMB has approved the information collection requirements contained in this part under control number 3150-0046. (b) The approved information collection requirements contained in this part appear in §§ 25.11, 25.17, 25.21, 25.23, 25.25, 25.27, 25.29, 25.31, 25.33, and 25.35. (c) This part contains information collection requirements in addition to those approved under the control number specified in paragraph (a) of this section. These information collection requirements and the control numbers under which they are approved are as follows: (1) In §§ 25.17(b), 25.21(c), 25.27(a), 25.29, and 25.31, NRC Form 237 is approved under control number 3150-0050. (2) In §§ 25.17(c), 25.21(c), 25.27(b), 25.29, and 25.31, the “Electronic Questionnaire for Investigations Processing (e-QIP), SF-86—Questionnaire for National Security Positions” is approved under control number 3206-0005. (3) In § 25.21(b), NRC Form 354 is approved under control number 3150-0026. (4) In § 25.33, NRC Form 136 is approved under control number 3150-0049. (5) In § 25.35, NRC Form 277 is approved under control number 3150-0051." 10:10:1.0.1.1.18.0.81.6,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.9 Communications.,NRC,,,"[68 FR 58803, Oct. 10, 2003, as amended at 74 FR 62681, Dec. 1, 2009; 80 FR 74979, Dec. 1, 2015]","Except where otherwise specified, communications and reports concerning the regulations in this part should be addressed to the Director, Division of Facilities and Security, Mail Stop T7-D57, and sent either by mail to the U.S. Nuclear Regulatory Commission, Washington, DC 20555-0001; by hand delivery to the NRC's offices at 11555 Rockville Pike, Rockville, Maryland; or, where practicable, by electronic submission, for example, Electronic Information Exchange, or CD-ROM. Electronic submissions must be made in a manner that enables the NRC to receive, read, authenticate, distribute, and archive the submission, and process and retrieve it a single page at a time. Detailed guidance on making electronic submissions can be obtained by visiting the NRC's Web site at http://www.nrc.gov/site-help/e-submittals.html; by e-mail to MSHD.Resource@nrc.gov; or by writing the Office of the Chief Information Officer, U.S. Nuclear Regulatory Commission, Washington, DC 20555-0001. The guidance discusses, among other topics, the formats the NRC can accept, the use of electronic signatures, and the treatment of nonpublic information." 10:10:1.0.1.1.18.0.81.7,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.11 Specific exemptions.,NRC,,,"[64 FR 15647, Apr. 1, 1999]","The NRC may, upon application by any interested person or upon its own initiative, grant exemptions from the requirements of the regulations of this part, that are— (a) Authorized by law, will not present an undue risk to the public health and safety, and are consistent with the common defense and security; or (b) Coincidental with one or more of the following: (1) An application of the regulation in the particular circumstances conflicts with other NRC rules or requirements; (2) An application of the regulation in the particular circumstances would not serve the underlying purpose of the rule or is not necessary to achieve the underlying purpose of the rule; (3) When compliance would result in undue hardship or other costs that significantly exceed those contemplated when the regulation was adopted, or that significantly exceed those incurred by others similarly situated; (4) When the exemption would result in benefit to the common defense and security that compensates for any decrease in the security that may result from the grant of the exemption; (5) When the exemption would provide only temporary relief from the applicable regulation and the licensee or applicant has made good faith efforts to comply with the regulation; (6) When there is any other material circumstance present that was not considered when the regulation was adopted that would be in the public interest to grant an exemption. If this condition is relied on exclusively for satisfying paragraph (b) of this section, the exemption may not be granted until the Executive Director for Operations has consulted with the Commission." 10:10:1.0.1.1.18.0.81.8,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.13 Maintenance of records.,NRC,,,"[45 FR 14481, Mar. 5, 1980, as amended at 53 FR 19245, May 27, 1988; 62 FR 17687, Apr. 11, 1997]","(a) Each licensee or organization employing individuals approved for personnel security access authorization under this part, shall maintain records as prescribed within the part. These records are subject to review and inspection by CSA representatives during security reviews. (b) Each record required by this part must be legible throughout the retention period specified by each Commission regulation. The record may be the original or a reproduced copy or a microform provided that the copy or microform is authenticated by authorized personnel and that the microform is capable of producing a clear copy throughout the required retention period. The record may also be stored in electronic media with the capability for producing legible, accurate, and complete records during the required retention period. Records such as letters, drawings, specifications, must include all pertinent information such as stamps, initials, and signatures. The licensee shall maintain adequate safeguards against tampering with and loss of records." 10:10:1.0.1.1.18.0.82.10,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.17 Approval for processing applicants for access authorization.,NRC,,,"[62 FR 17687, Apr. 11, 1997, as amended at 68 FR 62512, Nov. 5, 2003; 70 FR 32227, June 2, 2005; 72 FR 27411, May 16, 2007; 77 FR 26153, May 3, 2012; 77 FR 46258, Aug. 3, 2012; 86 FR 43401, Aug. 9, 2021; 86 FR 47209, Aug. 24, 2021; 87 FR 45241, July 28, 2022]","(a) Access authorizations must be requested for licensee employees or other persons (e.g., 10 CFR part 2, subpart I) who need access to classified information in connection with activities under 10 CFR parts 50, 52, 54, 60, 63, 70, 72, or 76. (b) The request must be submitted to the facility CSA. If the NRC is the CSA, the procedures in § 25.17 (c) and (d) will be followed. If the NRC is not the CSA, the request will be submitted to the CSA in accordance with procedures established by the CSA. The NRC will be notified of the request by a letter that includes the name, Social Security number and level of access authorization. (c) The request must include a completed personnel security packet (see § 25.17(d)) and request form (NRC Form 237) signed by a licensee, licensee contractor official, or other authorized person. (d)(1) Each personnel security packet submitted must include the following completed forms: (i) Electronic Questionnaire for Investigations Processing (e-QIP), SF-86 Questionnaire for National Security Positions; (ii) Two standard fingerprint cards (FD-258); (iii) Security Acknowledgment (NRC Form 176); and (iv) Other related forms where specified in accompanying instructions (NRC Form 254). (2) Only a Security Acknowledgment (NRC Form 176) need be completed by any person possessing an active access authorization, or who is being processed for an access authorization, by another Federal agency. The active or pending access authorization must be at an equivalent level to that required by the NRC and be based on an adequate investigation of not more than five years old. (e) To avoid delays in processing requests for access authorizations, each security packet should be reviewed for completeness and correctness (including legibility of response on the forms) before submittal. (f) The Defense Counterintelligence and Security Agency (DCSA) bills the NRC for the cost of each background investigation conducted in support of an application for access authorization (application). The combined cost of the DCSA investigation and the NRC's application processing overhead (NRC processing fee) are recovered through an access authorization fee imposed on applicants for access authorization. (1) Each application for access authorization, renewal, or change in level must be accompanied by a remittance, payable to the U.S. Nuclear Regulatory Commission, which is equal to the NRC access authorization fee. This fee must be determined using the following formula: the DCSA investigation billing rates on the day the NRC receives the application + the NRC processing fee = the NRC access authorization fee. The NRC processing fee is determined by multiplying the DCSA investigation billing rate on the day the NRC receives the application by 90.2 percent ( i.e., DCSA rate × 90.2 percent). (2) Updated DCSA investigation billing rates are published periodically in a Federal Investigations Notice (FIN) issued by the DCSA's Federal Investigative Services. Copies of the current DCSA investigation billing rates schedule can be obtained by contacting the NRC's Personnel Security Branch, Division of Facilities Security, Office of Administration by email to Licensee_ Access_ Authorization_ Fee.Resource@nrc.gov (3) The NRC's Information Access Authority Program (IAAP) is considered reimbursable work representing services provided to an organization for which the NRC is entitled payment. The NRC is authorized to receive and retain fees from licensees for services performed. The NRC's Office of the Chief Financial Officer periodically reviews the fees charged for IAAP and makes recommendations on revising those charges to reflect costs incurred by the NRC in providing those services. The reviews are performed using cost analysis techniques to determine the direct and indirect costs. Based on this review, the IAAP fees are adjusted to reflect the current cost for the program. IAAP requests for reciprocity will be charged a flat fee rate of $95.00 as referenced in paragraph (f)(4) of this section. This flat fee is aligned with the level of effort that has been expended by DCSA to process reciprocity requests, and accounts for inflation as well as recovery of the appropriate cost for conducting the investigations. Copies of the current NRC access authorization fee may be obtained by contacting the NRC's Personnel Security Branch, Division of Facilities and Security, Office of Administration by email at: Licensee_Access_ Authorization_ Fee.Resource@nrc.gov . Any change in the NRC's access authorization fee will be applicable to each access authorization request received on or after the effective date of the DCSA's most recently published investigation billing rates schedule. (4) Certain applications from individuals having current Federal access authorizations may be processed more expeditiously and at less cost because the Commission, at its discretion, may decide to accept the certification of access authorization and investigative data from other Federal Government agencies that grant personnel access authorizations. (i) Applications for reciprocity will be processed at the NRC flat fee rate of $95 per request, as referenced in the following table: 1 If the NRC determines, based on its review of available data, that a Tier 3 investigation is necessary, the appropriate NRC-L fee will be assessed as shown in appendix A to this part before the conduct of the investigation. 2 If the NRC determines, based on its review of available data, that a Tier 5 investigation is necessary, the appropriate NRC-Q fee will be assessed as shown in appendix A to this part before the conduct of the investigation. (ii) Applicants shall, in cases where reciprocity is not acceptable and it is necessary to perform a background investigation, be charged the appropriate fee referenced in appendix A to this part. Applicants shall calculate the access authorization fee according to the stated formula ( i.e., DCSA rate × 90.2 percent)." 10:10:1.0.1.1.18.0.82.11,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.19 Processing applications.,NRC,,,"[64 FR 15648, Apr. 1, 1999]","Each application for an access authorization or access authorization renewal must be submitted to the CSA. If the NRC is the CSA, the application and its accompanying fee must be submitted to the NRC Division of Facilities and Security. If necessary, the NRC Division of Facilities and Security may obtain approval from the appropriate Commission office exercising licensing or regulatory authority before processing the access authorization or access authorization renewal request. If the applicant is disapproved for processing, the NRC Division of Facilities and Security shall notify the submitter in writing and return the original application (security packet) and its accompanying fee." 10:10:1.0.1.1.18.0.82.12,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.21 Determination of initial and continued eligibility for access authorization.,NRC,,,"[62 FR 17688, Apr. 11, 1997, as amended at 64 FR 15648, Apr. 1, 1999]","(a) Following receipt by the CSA of the reports of the personnel security investigations, the record will be reviewed to determine that granting an access authorization or renewal of access authorization will not endanger the common defense and security and is clearly consistent with the national interest. If this determination is made, access authorization will be granted or renewed. If the NRC is the CSA, questions as to initial or continued eligibility will be determined in accordance with part 10 of chapter I. If another agency is the CSA, that agency will, under the requirements of the NISPOM, have established procedures at the facility to resolve questions as to initial or continued eligibility for access authorization. These questions will be determined in accordance with established CSA procedures already in effect for the facility. (b) The CSA must be promptly notified of developments that bear on continued eligibility for access authorization throughout the period for which the authorization is active (e.g., persons who marry subsequent to the completion of a personnel security packet must report this change by submitting a completed NRC Form 354, “Data Report on Spouse” or equivalent CSA form). (c)(1) Except as provided in paragraph (c)(2) of this section, an NRC “Q” access authorization must be renewed every five years from the date of issuance. Except as provided in paragraph (c)(2) of this section, an NRC “L” access authorization must be renewed every ten years from the date of issuance. An application for renewal must be submitted at least 120 days before the expiration of the five-year period for a “Q” access authorization and the ten-year period for an “L” access authorization, and must include: (i) A statement by the licensee or other person that the individual continues to require access to classified National Security Information or Restricted Data; and (ii) A personnel security packet as described in § 25.17(d). (2) Renewal applications and the required paperwork are not required for individuals who have a current and active access authorization from another Federal agency and who are subject to a reinvestigation program by that agency that is determined by the NRC to meet the NRC's requirements. (The DOE Reinvestigation Program has been determined to meet the NRC's requirements.) For these individuals, the submission of the SF-86 by the licensee or other person to the other Government agency pursuant to their reinvestigation requirements will satisfy the NRC's renewal submission and paperwork requirements, even if less than five years have passed since the date of issuance or renewal of the NRC “Q” access authorization, or if less than 10 years have passed since the date of issuance or renewal of the NRC “L” access authorization. Any NRC access authorization continued in response to the provisions of this paragraph will, thereafter, not be due for renewal until the date set by the other Government agency for the next reinvestigation of the individual pursuant to the other agency's reinvestigation program. However, the period of time for the initial and each subsequent NRC “Q” renewal application to the NRC may not exceed seven years or, in the case of an NRC “L” renewal application, twelve years. Any individual who is subject to the reinvestigation program requirements of another Federal agency but, for administrative or other reasons, does not submit reinvestigation forms to that agency within seven years for a “Q” renewal or twelve years for an “L” renewal of the previous submission, shall submit a renewal application to the NRC using the forms prescribed in § 25.17(d) before the expiration of the seven-year period for a “Q” renewal or twelve-year period for an “L” renewal. (3) If the NRC is not the CSA, reinvestigation program procedures and requirements will be set by the CSA." 10:10:1.0.1.1.18.0.82.13,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.23 Notification of grant of access authorization.,NRC,,,"[62 FR 17688, Apr. 11, 1997, as amended at 64 FR 15648, Apr. 1, 1999]","The determination to grant or renew access authorization will be furnished in writing to the licensee or organization that initiated the request. Upon receipt of the notification of original grant of access authorization, the licensee or organization shall obtain, as a condition for grant of access authorization and access to classified information, an executed “Classified Information Nondisclosure Agreement” (SF-312) from the affected individual. The SF-312 is an agreement between the United States and an individual who is cleared for access to classified information. An employee issued an initial access authorization shall execute a SF-312 before being granted access to classified information. The licensee or other organization shall forward the executed SF-312 to the CSA for retention. If the employee refuses to execute the SF-312, the licensee or other organization shall deny the employee access to classified information and submit a report to the CSA. The SF-312 must be signed and dated by the employee and witnessed. The employee's and witness' signatures must bear the same date. The individual shall also be given a security orientation briefing in accordance with § 95.33 of this chapter. Records of access authorization grant and renewal notification must be maintained by the licensee or other organization for three years after the access authorization has been terminated by the CSA. This information may also be furnished to other representatives of the Commission, to licensees, contractors, or other Federal agencies. Notifications of access authorization will not be given in writing to the affected individual except: (a) In those cases when the determination was made as a result of a Personnel Security Hearing or by a Personnel Security Review Panel ; or (b) When the individual also is the official designated by the licensee or other organization to whom written NRC notifications are forwarded." 10:10:1.0.1.1.18.0.82.14,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.25 Cancellation of requests for access authorization.,NRC,,,"[64 FR 15648, Apr. 1, 1999]","When a request for an individual's access authorization or renewal of an access authorization is withdrawn or canceled, the requestor shall notify the CSA immediately by telephone so that the single scope background investigation, national agency check with law and credit investigation, or other personnel security action may be discontinued. The requestor shall identify the full name and date of birth of the individual, the date of request, and the type of access authorization or access authorization renewal requested. The requestor shall confirm each telephone notification promptly in writing." 10:10:1.0.1.1.18.0.82.15,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.27 Reopening of cases in which requests for access authorizations are canceled.,NRC,,,"[62 FR 17689, Apr. 11, 1997, as amended at 64 FR 15648, Apr. 1, 1999]","(a) In conjunction with a new request for access authorization (NRC Form 237 or CSA equivalent) for individuals whose cases were previously canceled, new fingerprint cards (FD-257) in duplicate and a new Security Acknowledgment (NRC Form 176), or CSA equivalent, must be furnished to the CSA along with the request. (b) Additionally, if 90 days or more have elapsed since the date of the last Questionnaire for National Security Positions (SF-86), or CSA equivalent, the individual must complete a personnel security packet (see § 25.17(d)). The CSA, based on investigative or other needs, may require a complete personnel security packet in other cases as well. A fee, equal to the amount paid for an initial request, will be charged only if a new or updating investigation by the NRC is required." 10:10:1.0.1.1.18.0.82.16,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.29 Reinstatement of access authorization.,NRC,,,"[62 FR 17689, Apr. 11, 1997]","(a) An access authorization can be reinstated provided that: (1) No more than 24 months has lapsed since the date of termination of the clearance; (2) There has been no break in employment with the employer since the date of termination of the clearance; (3) There is no known adverse information; (4) The most recent investigation must not exceed 5 years (Top Secret, Q) or 10 years (Secret, L); and (5) The most recent investigation must meet or exceed the scope of the investigation required for the level of access authorization that is to be reinstated or granted. (b) An access authorization can be reinstated at the same, or lower, level by submission of a CSA-designated form to the CSA. The employee may not have access to classified information until receipt of written confirmation of reinstatement and an up-to-date personnel security packet will be furnished with the request for reinstatement of an access authorization. A new Security Acknowledgement will be obtained in all cases. Where personnel security packets are not required, a request for reinstatement must state the level of access authorization to be reinstated and the full name and date of birth of the individual to establish positive identification. A fee, equal to the amount paid for an initial request, will be charged only if a new or updating investigation by the NRC is required." 10:10:1.0.1.1.18.0.82.17,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.31 Extensions and transfers of access authorizations.,NRC,,,"[45 FR 14481, Mar. 5, 1980, as amended at 48 FR 24320, June 1, 1983; 57 FR 3721, Jan. 31, 1992; 62 FR 17689, Apr. 11, 1997; 64 FR 15648, Apr. 1, 1999]","(a) The NRC Division of Facilities and Security may, on request, extend the authorization of an individual who possesses an access authorization in connection with a particular employer or activity to permit access to classified information in connection with an assignment with another employer or activity. (b) The NRC Division of Facilities and Security may, on request, transfer an access authorization when an individual's access authorization under one employer or activity is terminated, simultaneously with the individual being granted an access authorization for another employer or activity. (c) Requests for an extension or transfer of an access authorization must state the full name of the person, date of birth, and level of access authorization. The Director, Division of Facilities and Security, may require a new personnel security packet (see § 25.17(c)) to be completed by the applicant. A fee, equal to the amount paid for an initial request, will be charged only if a new or updating investigation by the NRC is required. (d) The date of an extension or transfer of access authorization may not be used to determine when a request for renewal of access authorization is required. Access authorization renewal requests must be timely submitted, in accordance with § 25.21(c)." 10:10:1.0.1.1.18.0.82.18,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.33 Termination of access authorizations.,NRC,,,"[62 FR 17689, Apr. 11, 1997, as amended at 64 FR 15649, Apr. 1, 1999]","(a) Access authorizations will be terminated when: (1) An access authorization is no longer required; (2) An individual is separated from the employment or the activity for which he or she obtained an access authorization for a period of 90 days or more; or (3) An individual, pursuant to 10 CFR part 10 or other CSA-approved adjudicatory standards, is no longer eligible for an access authorization. (b) A representative of the licensee or other organization that employs the individual whose access authorization will be terminated shall immediately notify the CSA when the circumstances noted in paragraph (a)(1) or (a)(2) of this section exist; inform the individual that his or her access authorization is being terminated, and the reason; and that he or she will be considered for reinstatement of an access authorization if he or she resumes work requiring the authorization. (c) When an access authorization is to be terminated, a representative of the licensee or other organization shall conduct a security termination briefing of the individual involved, explain the Security Termination Statement (NRC Form 136 or CSA approved form) and have the individual complete the form. The representative shall promptly forward the original copy of the completed Security Termination Statement to CSA." 10:10:1.0.1.1.18.0.82.9,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.15 Access permitted under “Q” or “L” access authorization.,NRC,,,"[45 FR 14481, Mar. 5, 1980, as amended at 47 FR 9195, Mar. 4, 1982; 50 FR 36984, Sept. 11, 1985]","(a) A “Q” access authorization permits an individual access on a need-to-know basis to (1) Secret and Confidential Restricted Data and (2) Secret and Confidential National Security Information including intelligence information, CRYPTO ( i.e. , cryptographic information) or other classified communications security (COMSEC) information. (b) An “L” access authorization permits an individual access on a need-to-know basis to Confidential Restricted Data and Secret and Confidential National Security Information other than the categories specifically included in paragraph (a) of this section. In addition, access to certain Confidential COMSEC information is permitted as authorized by a National Communications Security Committee waiver dated February 14, 1985. (c) Each employee of the Commission is processed for one of the two levels of access authorization. Licensees and other persons will furnish National Security Information and/or Restricted Data to a Commission employee on official business when the employee has the appropriate level of NRC access authorization and need-to-know. Some individuals are permitted to begin NRC employment without an access authorization. However, no NRC employee shall be permitted access to any classified information until the appropriate level of access authorization has been granted to that employee by NRC." 10:10:1.0.1.1.18.0.83.19,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.35 Classified visits.,NRC,,,"[62 FR 17689, Apr. 11, 1997, as amended at 64 FR 15649, Apr. 1, 1999; 72 FR 49488, Aug. 28, 2007]","(a) The number of classified visits must be held to a minimum. The licensee, certificate holder, applicant for a standard design certification under part 52 of this chapter (including an applicant after the Commission has adopted a final standard design certification rule under part 52 of this chapter), or other facility, or an applicant for or holder of a standard design approval under part 52 of this chapter shall determine that the visit is necessary and that the purpose of the visit cannot be achieved without access to, or disclosure of, classified information. All classified visits require advance notification to, and approval of, the organization to be visited. In urgent cases, visit information may be furnished by telephone and confirmed in writing. (b) Representatives of the Federal Government, when acting in their official capacities as inspectors, investigators, or auditors, may visit a licensee, certificate holder, or other facility without furnishing advanced notification, provided these representatives present appropriate Government credentials upon arrival. Normally, however, Federal representatives will provide advance notification in the form of an NRC Form 277, “Request for Visit or Access Approval,” with the “need-to-know” certified by the appropriate NRC office exercising licensing or regulatory authority and verification of an NRC access authorization by the Division of Facilities and Security. (c) The licensee, certificate holder, or others shall include the following information on all Visit Authorization Letters (VAL) which they prepare. (1) Visitor's name, address, and telephone number and certification of the level of the facility security clearance; (2) Name, date and place of birth, and citizenship of the individual intending to visit; (3) Certification of the proposed visitor's personnel clearance and any special access authorizations required for the visit; (4) Name of person(s) to be visited; (5) Purpose and sufficient justification for the visit to allow for a determination of the necessity of the visit; and (6) Date or period during which the VAL is to be valid. (d) Classified visits may be arranged for a 12 month period. The requesting facility shall notify all places honoring these visit arrangements of any change in the individual's status that will cause the visit request to be canceled before its normal termination date. (e) The responsibility for determining need-to-know in connection with a classified visit rests with the individual who will disclose classified information during the visit. The licensee, certificate holder or other facility shall establish procedures to ensure positive identification of visitors before the disclosure of any classified information." 10:10:1.0.1.1.18.0.84.20,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.37 Violations.,NRC,,,"[48 FR 24320, June 1, 1983, as amended at 57 FR 55072, Nov. 24, 1992; 64 FR 15649, Apr. 1, 1999; 70 FR 32227, June 2, 2005; 75 FR 73941, Nov. 30, 2010]","(a) An injunction or other court order may be obtained to prohibit a violation of any provision of: (1) The Atomic Energy Act of 1954, as amended; (2) Title II of the Energy Reorganization Act of 1974, as amended; or (3) Any regulation or order issued under these Acts. (b) National Security Information is protected under the requirements and sanctions of Executive Order 13526, as amended, or any predecessor or successor orders." 10:10:1.0.1.1.18.0.84.21,10,Energy,I,,25,PART 25—ACCESS AUTHORIZATION,,,,§ 25.39 Criminal penalties.,NRC,,,"[57 FR 55072, Nov. 24, 1992]","(a) Section 223 of the Atomic Energy Act of 1954, as amended, provides for criminal sanctions for willful violation of, attempted violation of, or conspiracy to violate, any regulation issued under sections 161b, 161i, or 161o of the Act. For purposes of section 223, all the regulations in part 25 are issued under one or more of sections 161b, 161i, or 161o, except for the sections listed in paragraph (b) of this section. (b) The regulations in part 25 that are not issued under sections 161b, 161i, or 161o for the purposes of section 223 are as follows: §§ 25.1, 25.3, 25.5, 25.7, 25.8, 25.9, 25.11, 25.19, 25.25, 25.27, 25.29, 25.31, 25.37, and 25.39." 14:14:1.0.1.3.13.1.74.1,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,A,Subpart A—General,,§ 25.1 Applicability.,FAA,,,,"(a) This part prescribes airworthiness standards for the issue of type certificates, and changes to those certificates, for transport category airplanes. (b) Each person who applies under Part 21 for such a certificate or change must show compliance with the applicable requirements in this part." 14:14:1.0.1.3.13.1.74.2,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,A,Subpart A—General,,§ 25.2 Special retroactive requirements.,FAA,,,"[Amdt. 25-72, 55 FR 29773, July 20, 1990, as amended by Amdt. 25-99, 65 FR 36266, June 7, 2000]","The following special retroactive requirements are applicable to an airplane for which the regulations referenced in the type certificate predate the sections specified below— (a) Irrespective of the date of application, each applicant for a supplemental type certificate (or an amendment to a type certificate) involving an increase in passenger seating capacity to a total greater than that for which the airplane has been type certificated must show that the airplane concerned meets the requirements of: (1) Sections 25.721(d), 25.783(g), 25.785(c), 25.803(c)(2) through (9), 25.803 (d) and (e), 25.807 (a), (c), and (d), 25.809 (f) and (h), 25.811, 25.812, 25.813 (a), (b), and (c), 25.815, 25.817, 25.853 (a) and (b), 25.855(a), 25.993(f), and 25.1359(c) in effect on October 24, 1967, and (2) Sections 25.803(b) and 25.803(c)(1) in effect on April 23, 1969. (b) Irrespective of the date of application, each applicant for a supplemental type certificate (or an amendment to a type certificate) for an airplane manufactured after October 16, 1987, must show that the airplane meets the requirements of § 25.807(c)(7) in effect on July 24, 1989. (c) Compliance with subsequent revisions to the sections specified in paragraph (a) or (b) of this section may be elected or may be required in accordance with § 21.101(a) of this chapter." 14:14:1.0.1.3.13.1.74.3,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,A,Subpart A—General,,§ 25.3 Special provisions for ETOPS type design approvals.,FAA,,,"[Doc. No. FAA-2002-6717, 72 FR 1873, Jan. 16, 2007]","(a) Applicability. This section applies to an applicant for ETOPS type design approval of an airplane: (1) That has an existing type certificate on February 15, 2007; or (2) For which an application for an original type certificate was submitted before February 15, 2007. (b) Airplanes with two engines. (1) For ETOPS type design approval of an airplane up to and including 180 minutes, an applicant must comply with § 25.1535, except that it need not comply with the following provisions of Appendix K, K25.1.4, of this part: (i) K25.1.4(a), fuel system pressure and flow requirements; (ii) K25.1.4(a)(3), low fuel alerting; and (iii) K25.1.4(c), engine oil tank design. (2) For ETOPS type design approval of an airplane beyond 180 minutes an applicant must comply with § 25.1535. (c) Airplanes with more than two engines. An applicant for ETOPS type design approval must comply with § 25.1535 for an airplane manufactured on or after February 17, 2015, except that, for an airplane configured for a three person flight crew, the applicant need not comply with Appendix K, K25.1.4(a)(3), of this part, low fuel alerting." 14:14:1.0.1.3.13.1.74.4,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,A,Subpart A—General,,§ 25.4 Definitions.,FAA,,,"[Doc. No. FAA-2022-1544, 89 FR 68731, Aug. 27, 2024]","(a) For the purposes of this part, the following general definitions apply: (1) Certification maintenance requirement means a required scheduled maintenance task established during the design certification of the airplane systems as an airworthiness limitation of the type certificate or supplemental type certificate. (2) Significant latent failure is a latent failure that, in combination with one or more specific failures or events, would result in a hazardous or catastrophic failure condition. (b) For purposes of this part, the following failure conditions, in order of increasing severity, apply: (1) Major failure condition means a failure condition that would reduce the capability of the airplane or the ability of the flightcrew to cope with adverse operating conditions, to the extent that there would be— (i) A significant reduction in safety margins or functional capabilities, (ii) A physical discomfort or a significant increase in flightcrew workload or in conditions impairing the efficiency of the flightcrew, (iii) Physical distress to passengers or cabin crew, possibly including injuries, or (iv) An effect of similar severity. (2) Hazardous failure condition means a failure condition that would reduce the capability of the airplane or the ability of the flightcrew to cope with adverse operating conditions, to the extent that there would be— (i) A large reduction in safety margins or functional capabilities, (ii) Physical distress or excessive workload such that the flightcrew cannot be relied upon to perform their tasks accurately or completely, or (iii) Serious or fatal injuries to a relatively small number of persons other than the flightcrew. (3) Catastrophic failure condition means a failure condition that would result in multiple fatalities, usually with the loss of the airplane. (c) For purposes of this part, the following failure conditions in order of decreasing probability apply: (1) Probable failure condition means a failure condition that is anticipated to occur one or more times during the entire operational life of each airplane of a given type. (2) Remote failure condition means a failure condition that is not anticipated to occur to each airplane of a given type during its entire operational life, but which may occur several times during the total operational life of a number of airplanes of a given type. (3) Extremely remote failure condition means a failure condition that is not anticipated to occur to each airplane of a given type during its entire operational life, but which may occur a few times during the total operational life of all airplanes of a given type. (4) Extremely improbable failure condition means a failure condition that is not anticipated to occur during the total operational life of all airplanes of a given type." 14:14:1.0.1.3.13.1.74.5,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,A,Subpart A—General,,§ 25.5 Incorporations by reference.,FAA,,,"[73 FR 42494, July 21, 2008, as amended by Doc. No. FAA-2018-0119, Amdt. 21-101, 83 FR 9169, Mar. 5, 2018]","(a) The materials listed in this section are incorporated by reference in the corresponding sections noted. These incorporations by reference were approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. These materials are incorporated as they exist on the date of the approval, and notice of any change in these materials will be published in the Federal Register. The materials are available for purchase at the corresponding addresses noted below, and all are available for inspection at the National Archives and Records Administration (NARA). For information on the availability of this material at NARA, call 202-741-6030, or go to: http://www.archives.gov/federal-register/cfr/ibr-locations.html. (b) The following materials are available for purchase from the following address: The National Technical Information Services (NTIS), Springfield, Virginia 22166. (1) Fuel Tank Flammability Assessment Method User's Manual, dated May 2008, document number DOT/FAA/AR-05/8, IBR approved for § 25.981 and Appendix N. It can also be obtained at the following Web site: http://www.fire.tc.faa.gov/systems/fueltank/FTFAM.stm. (2) [Reserved]" 14:14:1.0.1.3.13.2.74.1,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.21 Proof of compliance.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990; Amdt. 25-121, 72 FR 44665, Aug. 8, 2007 Amdt. 25-135, 76 FR 74654, Dec. 1, 2011; Amdt. 25-140, 79 FR 65524, Nov. 4, 2014]","(a) Each requirement of this subpart must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested. This must be shown— (1) By tests upon an airplane of the type for which certification is requested, or by calculations based on, and equal in accuracy to, the results of testing; and (2) By systematic investigation of each probable combination of weight and center of gravity, if compliance cannot be reasonably inferred from combinations investigated. (b) [Reserved] (c) The controllability, stability, trim, and stalling characteristics of the airplane must be shown for each altitude up to the maximum expected in operation. (d) Parameters critical for the test being conducted, such as weight, loading (center of gravity and inertia), airspeed, power, and wind, must be maintained within acceptable tolerances of the critical values during flight testing. (e) If compliance with the flight characteristics requirements is dependent upon a stability augmentation system or upon any other automatic or power-operated system, compliance must be shown with §§ 25.671 and 25.672. (f) In meeting the requirements of §§ 25.105(d), 25.125, 25.233, and 25.237, the wind velocity must be measured at a height of 10 meters above the surface, or corrected for the difference between the height at which the wind velocity is measured and the 10-meter height. (g) The requirements of this subpart associated with icing conditions apply only if the applicant is seeking certification for flight in icing conditions. (1) Paragraphs (g)(3) and (4) of this section apply only to airplanes with one or both of the following attributes: (i) Maximum takeoff gross weight is less than 60,000 lbs; or (ii) The airplane is equipped with reversible flight controls. (2) Each requirement of this subpart, except §§ 25.121(a), 25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 25.251(b) through (e), must be met in the icing conditions specified in Appendix C of this part. Section 25.207(c) and (d) must be met in the landing configuration in the icing conditions specified in Appendix C, but need not be met for other configurations. Compliance must be shown using the ice accretions defined in part II of Appendix C of this part, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual. (3) If the applicant does not seek certification for flight in all icing conditions defined in Appendix O of this part, each requirement of this subpart, except §§ 25.105, 25.107, 25.109, 25.111, 25.113, 25.115, 25.121, 25.123, 25.143(b)(1), (b)(2), and (c)(1), 25.149, 25.201(c)(2), 25.207(c), (d), and (e)(1), 25.239, and 25.251(b) through (e), must be met in the Appendix O icing conditions for which certification is not sought in order to allow a safe exit from those conditions. Compliance must be shown using the ice accretions defined in part II, paragraphs (b) and (d) of Appendix O, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual. (4) If the applicant seeks certification for flight in any portion of the icing conditions of Appendix O of this part, each requirement of this subpart, except §§ 25.121(a), 25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 25.251(b) through (e), must be met in the Appendix O icing conditions for which certification is sought. Section 25.207(c) and (d) must be met in the landing configuration in the Appendix O icing conditions for which certification is sought, but need not be met for other configurations. Compliance must be shown using the ice accretions defined in part II, paragraphs (c) and (d) of Appendix O, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual." 14:14:1.0.1.3.13.2.74.2,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.23 Load distribution limits.,FAA,,,,"(a) Ranges of weights and centers of gravity within which the airplane may be safely operated must be established. If a weight and center of gravity combination is allowable only within certain load distribution limits (such as spanwise) that could be inadvertently exceeded, these limits and the corresponding weight and center of gravity combinations must be established. (b) The load distribution limits may not exceed— (1) The selected limits; (2) The limits at which the structure is proven; or (3) The limits at which compliance with each applicable flight requirement of this subpart is shown." 14:14:1.0.1.3.13.2.74.3,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.25 Weight limits.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-63, 53 FR 16365, May 6, 1988]","(a) Maximum weights. Maximum weights corresponding to the airplane operating conditions (such as ramp, ground or water taxi, takeoff, en route, and landing), environmental conditions (such as altitude and temperature), and loading conditions (such as zero fuel weight, center of gravity position and weight distribution) must be established so that they are not more than— (1) The highest weight selected by the applicant for the particular conditions; or (2) The highest weight at which compliance with each applicable structural loading and flight requirement is shown, except that for airplanes equipped with standby power rocket engines the maximum weight must not be more than the highest weight established in accordance with appendix E of this part; or (3) The highest weight at which compliance is shown with the certification requirements of Part 36 of this chapter. (b) Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is not less than— (1) The lowest weight selected by the applicant; (2) The design minimum weight (the lowest weight at which compliance with each structural loading condition of this part is shown); or (3) The lowest weight at which compliance with each applicable flight requirement is shown." 14:14:1.0.1.3.13.2.74.4,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.27 Center of gravity limits.,FAA,,,,"The extreme forward and the extreme aft center of gravity limitations must be established for each practicably separable operating condition. No such limit may lie beyond— (a) The extremes selected by the applicant; (b) The extremes within which the structure is proven; or (c) The extremes within which compliance with each applicable flight requirement is shown." 14:14:1.0.1.3.13.2.74.5,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.29 Empty weight and corresponding center of gravity.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990]","(a) The empty weight and corresponding center of gravity must be determined by weighing the airplane with— (1) Fixed ballast; (2) Unusable fuel determined under § 25.959; and (3) Full operating fluids, including— (i) Oil; (ii) Hydraulic fluid; and (iii) Other fluids required for normal operation of airplane systems, except potable water, lavatory precharge water, and fluids intended for injection in the engine. (b) The condition of the airplane at the time of determining empty weight must be one that is well defined and can be easily repeated." 14:14:1.0.1.3.13.2.74.6,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.31 Removable ballast.,FAA,,,,Removable ballast may be used on showing compliance with the flight requirements of this subpart. 14:14:1.0.1.3.13.2.74.7,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.33 Propeller speed and pitch limits.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29774, July 20, 1990]","(a) The propeller speed and pitch must be limited to values that will ensure— (1) Safe operation under normal operating conditions; and (2) Compliance with the performance requirements of §§ 25.101 through 25.125. (b) There must be a propeller speed limiting means at the governor. It must limit the maximum possible governed engine speed to a value not exceeding the maximum allowable r.p.m. (c) The means used to limit the low pitch position of the propeller blades must be set so that the engine does not exceed 103 percent of the maximum allowable engine rpm or 99 percent of an approved maximum overspeed, whichever is greater, with— (1) The propeller blades at the low pitch limit and governor inoperative; (2) The airplane stationary under standard atmospheric conditions with no wind; and (3) The engines operating at the takeoff manifold pressure limit for reciprocating engine powered airplanes or the maximum takeoff torque limit for turbopropeller engine-powered airplanes." 14:14:1.0.1.3.13.2.75.10,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.105 Takeoff.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-92, 63 FR 8318, Feb. 18, 1998; Amdt. 25-121, 72 FR 44665, Aug. 8, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) The takeoff speeds prescribed by § 25.107, the accelerate-stop distance prescribed by § 25.109, the takeoff path prescribed by § 25.111, the takeoff distance and takeoff run prescribed by § 25.113, and the net takeoff flight path prescribed by § 25.115, must be determined in the selected configuration for takeoff at each weight, altitude, and ambient temperature within the operational limits selected by the applicant— (1) In non-icing conditions; and (2) In icing conditions, if in the configuration used to show compliance with § 25.121(b), and with the most critical of the takeoff ice accretion(s) defined in appendices C and O of this part, as applicable, in accordance with § 25.21(g): (i) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or (ii) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b). (b) No takeoff made to determine the data required by this section may require exceptional piloting skill or alertness. (c) The takeoff data must be based on— (1) In the case of land planes and amphibians: (i) Smooth, dry and wet, hard-surfaced runways; and (ii) At the option of the applicant, grooved or porous friction course wet, hard-surfaced runways. (2) Smooth water, in the case of seaplanes and amphibians; and (3) Smooth, dry snow, in the case of skiplanes. (d) The takeoff data must include, within the established operational limits of the airplane, the following operational correction factors: (1) Not more than 50 percent of nominal wind components along the takeoff path opposite to the direction of takeoff, and not less than 150 percent of nominal wind components along the takeoff path in the direction of takeoff. (2) Effective runway gradients." 14:14:1.0.1.3.13.2.75.11,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.107 Takeoff speeds.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; Amdt. 25-92, 63 FR 8318, Feb. 18, 1998; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-121, 72 FR 44665, Aug. 8, 2007; Amdt. 25-135, 76 FR 74654, Dec. 1, 2011]","(a) V 1 must be established in relation to V EF as follows: (1) V EF is the calibrated airspeed at which the critical engine is assumed to fail. V EF must be selected by the applicant, but may not be less than V MCG determined under § 25.149(e). (2) V 1 , in terms of calibrated airspeed, is selected by the applicant; however, V 1 may not be less than V EF plus the speed gained with critical engine inoperative during the time interval between the instant at which the critical engine is failed, and the instant at which the pilot recognizes and reacts to the engine failure, as indicated by the pilot's initiation of the first action (e.g., applying brakes, reducing thrust, deploying speed brakes) to stop the airplane during accelerate-stop tests. (b) V 2MIN, in terms of calibrated airspeed, may not be less than— (1) 1.13 V SR for— (i) Two-engine and three-engine turbopropeller and reciprocating engine powered airplanes; and (ii) Turbojet powered airplanes without provisions for obtaining a significant reduction in the one-engine-inoperative power-on stall speed; (2) 1.08 V SR for— (i) Turbopropeller and reciprocating engine powered airplanes with more than three engines; and (ii) Turbojet powered airplanes with provisions for obtaining a significant reduction in the one-engine-inoperative power-on stall speed; and (3) 1.10 times V MC established under § 25.149. (c) V 2 , in terms of calibrated airspeed, must be selected by the applicant to provide at least the gradient of climb required by § 25.121(b) but may not be less than— (1) V 2MIN ; (2) V R plus the speed increment attained (in accordance with § 25.111(c)(2)) before reaching a height of 35 feet above the takeoff surface; and (3) A speed that provides the maneuvering capability specified in § 25.143(h). (d) V MU is the calibrated airspeed at and above which the airplane can safely lift off the ground, and con- tinue the takeoff. V MU speeds must be selected by the applicant throughout the range of thrust-to-weight ratios to be certificated. These speeds may be established from free air data if these data are verified by ground takeoff tests. (e) V R, in terms of calibrated airspeed, must be selected in accordance with the conditions of paragraphs (e)(1) through (4) of this section: (1) V R may not be less than— (i) V 1 ; (ii) 105 percent of V MC ; (iii) The speed (determined in accordance with § 25.111(c)(2)) that allows reaching V 2 before reaching a height of 35 feet above the takeoff surface; or (iv) A speed that, if the airplane is rotated at its maximum practicable rate, will result in a V LOF of not less than — (A) 110 percent of V MU in the all-engines-operating condition, and 105 percent of V MU determined at the thrust-to-weight ratio corresponding to the one-engine-inoperative condition; or (B) If the V MU attitude is limited by the geometry of the airplane ( i.e., tail contact with the runway), 108 percent of V MU in the all-engines-operating condition, and 104 percent of V MU determined at the thrust-to-weight ratio corresponding to the one-engine-inoperative condition. (2) For any given set of conditions (such as weight, configuration, and temperature), a single value of V R, obtained in accordance with this paragraph, must be used to show compliance with both the one-engine-inoperative and the all-engines-operating takeoff provisions. (3) It must be shown that the one-engine-inoperative takeoff distance, using a rotation speed of 5 knots less than V R established in accordance with paragraphs (e)(1) and (2) of this section, does not exceed the corresponding one-engine-inoperative takeoff distance using the established V R . The takeoff distances must be determined in accordance with § 25.113(a)(1). (4) Reasonably expected variations in service from the established takeoff procedures for the operation of the airplane (such as over-rotation of the airplane and out-of-trim conditions) may not result in unsafe flight characteristics or in marked increases in the scheduled takeoff distances established in accordance with § 25.113(a). (f) V LOF is the calibrated airspeed at which the airplane first becomes airborne. (g) V FTO , in terms of calibrated airspeed, must be selected by the applicant to provide at least the gradient of climb required by § 25.121(c), but may not be less than— (1) 1.18 V SR ; and (2) A speed that provides the maneuvering capability specified in § 25.143(h). (h) In determining the takeoff speeds V 1 , V R , and V 2 for flight in icing conditions, the values of V MCG , V MC , and V MU determined for non-icing conditions may be used." 14:14:1.0.1.3.13.2.75.12,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.109 Accelerate-stop distance.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-92, 63 FR 8318, Feb. 18, 1998]","(a) The accelerate-stop distance on a dry runway is the greater of the following distances: (1) The sum of the distances necessary to— (i) Accelerate the airplane from a standing start with all engines operating to V EF for takeoff from a dry runway; (ii) Allow the airplane to accelerate from V EF to the highest speed reached during the rejected takeoff, assuming the critical engine fails at V EF and the pilot takes the first action to reject the takeoff at the V 1 for takeoff from a dry runway; and (iii) Come to a full stop on a dry runway from the speed reached as prescribed in paragraph (a)(1)(ii) of this section; plus (iv) A distance equivalent to 2 seconds at the V 1 for takeoff from a dry runway. (2) The sum of the distances necessary to— (i) Accelerate the airplane from a standing start with all engines operating to the highest speed reached during the rejected takeoff, assuming the pilot takes the first action to reject the takeoff at the V 1 for takeoff from a dry runway; and (ii) With all engines still operating, come to a full stop on dry runway from the speed reached as prescribed in paragraph (a)(2)(i) of this section; plus (iii) A distance equivalent to 2 seconds at the V 1 for takeoff from a dry runway. (b) The accelerate-stop distance on a wet runway is the greater of the following distances: (1) The accelerate-stop distance on a dry runway determined in accordance with paragraph (a) of this section; or (2) The accelerate-stop distance determined in accordance with paragraph (a) of this section, except that the runway is wet and the corresponding wet runway values of V EF and V 1 are used. In determining the wet runway accelerate-stop distance, the stopping force from the wheel brakes may never exceed: (i) The wheel brakes stopping force determined in meeting the requirements of § 25.101(i) and paragraph (a) of this section; and (ii) The force resulting from the wet runway braking coefficient of friction determined in accordance with paragraphs (c) or (d) of this section, as applicable, taking into account the distribution of the normal load between braked and unbraked wheels at the most adverse center-of-gravity position approved for takeoff. (c) The wet runway braking coefficient of friction for a smooth wet runway is defined as a curve of friction coefficient versus ground speed and must be computed as follows: (1) The maximum tire-to-ground wet runway braking coefficient of friction is defined as: Where— Tire Pressure = maximum airplane operating tire pressure (psi); μ t/gMAX = maximum tire-to-ground braking coefficient; V = airplane true ground speed (knots); and Linear interpolation may be used for tire pressures other than those listed. Where— Tire Pressure = maximum airplane operating tire pressure (psi); μ t/gMAX = maximum tire-to-ground braking coefficient; V = airplane true ground speed (knots); and Linear interpolation may be used for tire pressures other than those listed. (2) The maximum tire-to-ground wet runway braking coefficient of friction must be adjusted to take into account the efficiency of the anti-skid system on a wet runway. Anti-skid system operation must be demonstrated by flight testing on a smooth wet runway, and its efficiency must be determined. Unless a specific anti-skid system efficiency is determined from a quantitative analysis of the flight testing on a smooth wet runway, the maximum tire-to-ground wet runway braking coefficient of friction determined in paragraph (c)(1) of this section must be multiplied by the efficiency value associated with the type of anti-skid system installed on the airplane: (d) At the option of the applicant, a higher wet runway braking coefficient of friction may be used for runway surfaces that have been grooved or treated with a porous friction course material. For grooved and porous friction course runways, the wet runway braking coefficent of friction is defined as either: (1) 70 percent of the dry runway braking coefficient of friction used to determine the dry runway accelerate-stop distance; or (2) The wet runway braking coefficient defined in paragraph (c) of this section, except that a specific anti-skid system efficiency, if determined, is appropriate for a grooved or porous friction course wet runway, and the maximum tire-to-ground wet runway braking coefficient of friction is defined as: Where— Tire Pressure = maximum airplane operating tire pressure (psi); μ t/gMAX = maximum tire-to-ground braking coefficient; V = airplane true ground speed (knots); and Linear interpolation may be used for tire pressures other than those listed. Where— Tire Pressure = maximum airplane operating tire pressure (psi); μ t/gMAX = maximum tire-to-ground braking coefficient; V = airplane true ground speed (knots); and Linear interpolation may be used for tire pressures other than those listed. (e) Except as provided in paragraph (f)(1) of this section, means other than wheel brakes may be used to determine the accelerate-stop distance if that means— (1) Is safe and reliable; (2) Is used so that consistent results can be expected under normal operating conditions; and (3) Is such that exceptional skill is not required to control the airplane. (f) The effects of available reverse thrust— (1) Shall not be included as an additional means of deceleration when determining the accelerate-stop distance on a dry runway; and (2) May be included as an additional means of deceleration using recommended reverse thrust procedures when determining the accelerate-stop distance on a wet runway, provided the requirements of paragraph (e) of this section are met. (g) The landing gear must remain extended throughout the accelerate-stop distance. (h) If the accelerate-stop distance includes a stopway with surface characteristics substantially different from those of the runway, the takeoff data must include operational correction factors for the accelerate-stop distance. The correction factors must account for the particular surface characteristics of the stopway and the variations in these characteristics with seasonal weather conditions (such as temperature, rain, snow, and ice) within the established operational limits. (i) A flight test demonstration of the maximum brake kinetic energy accelerate-stop distance must be conducted with not more than 10 percent of the allowable brake wear range remaining on each of the airplane wheel brakes." 14:14:1.0.1.3.13.2.75.13,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.111 Takeoff path.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-6, 30 FR 8468, July 2, 1965; Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55 FR 29774, July 20, 1990; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 2004; Amdt. 25-121, 72 FR 44666; Aug. 8, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) The takeoff path extends from a standing start to a point in the takeoff at which the airplane is 1,500 feet above the takeoff surface, or at which the transition from the takeoff to the en route configuration is completed and V FTO is reached, whichever point is higher. In addition— (1) The takeoff path must be based on the procedures prescribed in § 25.101(f); (2) The airplane must be accelerated on the ground to V EF, at which point the critical engine must be made inoperative and remain inoperative for the rest of the takeoff; and (3) After reaching V EF, the airplane must be accelerated to V 2 . (b) During the acceleration to speed V 2 , the nose gear may be raised off the ground at a speed not less than V R . However, landing gear retraction may not be begun until the airplane is airborne. (c) During the takeoff path determination in accordance with paragraphs (a) and (b) of this section— (1) The slope of the airborne part of the takeoff path must be positive at each point; (2) The airplane must reach V 2 before it is 35 feet above the takeoff surface and must continue at a speed as close as practical to, but not less than V 2 , until it is 400 feet above the takeoff surface; (3) At each point along the takeoff path, starting at the point at which the airplane reaches 400 feet above the takeoff surface, the available gradient of climb may not be less than— (i) 1.2 percent for two-engine airplanes; (ii) 1.5 percent for three-engine airplanes; and (iii) 1.7 percent for four-engine airplanes. (4) The airplane configuration may not be changed, except for gear retraction and automatic propeller feathering, and no change in power or thrust that requires action by the pilot may be made until the airplane is 400 feet above the takeoff surface; and (5) If § 25.105(a)(2) requires the takeoff path to be determined for flight in icing conditions, the airborne part of the takeoff must be based on the airplane drag: (i) With the most critical of the takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), from a height of 35 feet above the takeoff surface up to the point where the airplane is 400 feet above the takeoff surface; and (ii) With the most critical of the final takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), from the point where the airplane is 400 feet above the takeoff surface to the end of the takeoff path. (d) The takeoff path must be determined by a continuous demonstrated takeoff or by synthesis from segments. If the takeoff path is determined by the segmental method— (1) The segments must be clearly defined and must be related to the distinct changes in the configuration, power or thrust, and speed; (2) The weight of the airplane, the configuration, and the power or thrust must be constant throughout each segment and must correspond to the most critical condition prevailing in the segment; (3) The flight path must be based on the airplane's performance without ground effect; and (4) The takeoff path data must be checked by continuous demonstrated takeoffs up to the point at which the airplane is out of ground effect and its speed is stabilized, to ensure that the path is conservative relative to the continous path. The airplane is considered to be out of the ground effect when it reaches a height equal to its wing span. (e) For airplanes equipped with standby power rocket engines, the takeoff path may be determined in accordance with section II of appendix E." 14:14:1.0.1.3.13.2.75.14,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.113 Takeoff distance and takeoff run.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-92, 63 FR 8320, Feb. 18, 1998]","(a) Takeoff distance on a dry runway is the greater of— (1) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 35 feet above the takeoff surface, determined under § 25.111 for a dry runway; or (2) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to the point at which the airplane is 35 feet above the takeoff surface, as determined by a procedure consistent with § 25.111. (b) Takeoff distance on a wet runway is the greater of— (1) The takeoff distance on a dry runway determined in accordance with paragraph (a) of this section; or (2) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface, achieved in a manner consistent with the achievement of V 2 before reaching 35 feet above the takeoff surface, determined under § 25.111 for a wet runway. (c) If the takeoff distance does not include a clearway, the takeoff run is equal to the takeoff distance. If the takeoff distance includes a clearway— (1) The takeoff run on a dry runway is the greater of— (i) The horizontal distance along the takeoff path from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, as determined under § 25.111 for a dry runway; or (ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, determined by a procedure consistent with § 25.111. (2) The takeoff run on a wet runway is the greater of— (i) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface, achieved in a manner consistent with the achievement of V 2 before reaching 35 feet above the takeoff surface, as determined under § 25.111 for a wet runway; or (ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, determined by a procedure consistent with § 25.111." 14:14:1.0.1.3.13.2.75.15,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.115 Takeoff flight path.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-92, 63 FR 8320, Feb. 18, 1998]","(a) The takeoff flight path shall be considered to begin 35 feet above the takeoff surface at the end of the takeoff distance determined in accordance with § 25.113(a) or (b), as appropriate for the runway surface condition. (b) The net takeoff flight path data must be determined so that they represent the actual takeoff flight paths (determined in accordance with § 25.111 and with paragraph (a) of this section) reduced at each point by a gradient of climb equal to— (1) 0.8 percent for two-engine airplanes; (2) 0.9 percent for three-engine airplanes; and (3) 1.0 percent for four-engine airplanes. (c) The prescribed reduction in climb gradient may be applied as an equivalent reduction in acceleration along that part of the takeoff flight path at which the airplane is accelerated in level flight." 14:14:1.0.1.3.13.2.75.16,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.117 Climb: general.,FAA,,,,"Compliance with the requirements of §§ 25.119 and 25.121 must be shown at each weight, altitude, and ambient temperature within the operational limits established for the airplane and with the most unfavorable center of gravity for each configuration." 14:14:1.0.1.3.13.2.75.17,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.119 Landing climb: All-engines-operating.,FAA,,,"[Amdt. 25-121, 72 FR 44666; Aug. 8, 2007, as amended by Amdt. 25-,140, 79 FR 65525, Nov. 4, 2014]","In the landing configuration, the steady gradient of climb may not be less than 3.2 percent, with the engines at the power or thrust that is available 8 seconds after initiation of movement of the power or thrust controls from the minimum flight idle to the go-around power or thrust setting— (a) In non-icing conditions, with a climb speed of V REF determined in accordance with § 25.125(b)(2)(i); and (b) In icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), and with a climb speed of V REF determined in accordance with § 25.125(b)(2)(ii)." 14:14:1.0.1.3.13.2.75.18,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.121 Climb: One-engine-inoperative.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-121, 72 FR 44666; Aug. 8, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) Takeoff; landing gear extended. In the critical takeoff configuration existing along the flight path (between the points at which the airplane reaches V LOF and at which the landing gear is fully retracted) and in the configuration used in § 25.111 but without ground effect, the steady gradient of climb must be positive for two-engine airplanes, and not less than 0.3 percent for three-engine airplanes or 0.5 percent for four-engine airplanes, at V LOF and with— (1) The critical engine inoperative and the remaining engines at the power or thrust available when retraction of the landing gear is begun in accordance with § 25.111 unless there is a more critical power operating condition existing later along the flight path but before the point at which the landing gear is fully retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun, determined under § 25.111. (b) Takeoff; landing gear retracted. In the takeoff configuration existing at the point of the flight path at which the landing gear is fully retracted, and in the configuration used in § 25.111 but without ground effect: (1) The steady gradient of climb may not be less than 2.4 percent for two-engine airplanes, 2.7 percent for three-engine airplanes, and 3.0 percent for four-engine airplanes, at V 2 with: (i) The critical engine inoperative, the remaining engines at the takeoff power or thrust available at the time the landing gear is fully retracted, determined under § 25.111, unless there is a more critical power operating condition existing later along the flight path but before the point where the airplane reaches a height of 400 feet above the takeoff surface; and (ii) The weight equal to the weight existing when the airplane's landing gear is fully retracted, determined under § 25.111. (2) The requirements of paragraph (b)(1) of this section must be met: (i) In non-icing conditions; and (ii) In icing conditions with the most critical of the takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if in the configuration used to show compliance with § 25.121(b) with this takeoff ice accretion: (A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or (B) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b). (c) Final takeoff. In the en route configuration at the end of the takeoff path determined in accordance with § 25.111: (1) The steady gradient of climb may not be less than 1.2 percent for two-engine airplanes, 1.5 percent for three-engine airplanes, and 1.7 percent for four-engine airplanes, at V FTO with— (i) The critical engine inoperative and the remaining engines at the available maximum continuous power or thrust; and (ii) The weight equal to the weight existing at the end of the takeoff path, determined under § 25.111. (2) The requirements of paragraph (c)(1) of this section must be met: (i) In non-icing conditions; and (ii) In icing conditions with the most critical of the final takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if in the configuration used to show compliance with § 25.121(b) with the takeoff ice accretion used to show compliance with § 25.111(c)(5)(i): (A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or (B) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b). (d) Approach. In a configuration corresponding to the normal all-engines-operating procedure in which V SR for this configuration does not exceed 110 percent of the V SR for the related all-engines-operating landing configuration: (1) The steady gradient of climb may not be less than 2.1 percent for two-engine airplanes, 2.4 percent for three-engine airplanes, and 2.7 percent for four-engine airplanes, with— (i) The critical engine inoperative, the remaining engines at the go-around power or thrust setting; (ii) The maximum landing weight; (iii) A climb speed established in connection with normal landing procedures, but not exceeding 1.4 V SR ; and (iv) Landing gear retracted. (2) The requirements of paragraph (d)(1) of this section must be met: (i) In non-icing conditions; and (ii) In icing conditions with the most critical of the approach ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g). The climb speed selected for non-icing conditions may be used if the climb speed for icing conditions, computed in accordance with paragraph (d)(1)(iii) of this section, does not exceed that for non-icing conditions by more than the greater of 3 knots CAS or 3 percent." 14:14:1.0.1.3.13.2.75.19,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.123 En route flight paths.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-121, 72 FR 44666; Aug. 8, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) For the en route configuration, the flight paths prescribed in paragraph (b) and (c) of this section must be determined at each weight, altitude, and ambient temperature, within the operating limits established for the airplane. The variation of weight along the flight path, accounting for the progressive consumption of fuel and oil by the operating engines, may be included in the computation. The flight paths must be determined at a speed not less than V FTO , with— (1) The most unfavorable center of gravity; (2) The critical engines inoperative; (3) The remaining engines at the available maximum continuous power or thrust; and (4) The means for controlling the engine-cooling air supply in the position that provides adequate cooling in the hot-day condition. (b) The one-engine-inoperative net flight path data must represent the actual climb performance diminished by a gradient of climb of 1.1 percent for two-engine airplanes, 1.4 percent for three-engine airplanes, and 1.6 percent for four-engine airplanes— (1) In non-icing conditions; and (2) In icing conditions with the most critical of the en route ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if: (i) A speed of 1.18 “V SR0 with the en route ice accretion exceeds the en route speed selected for non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or (ii) The degradation of the gradient of climb is greater than one-half of the applicable actual-to-net flight path reduction defined in paragraph (b) of this section. (c) For three- or four-engine airplanes, the two-engine-inoperative net flight path data must represent the actual climb performance diminished by a gradient of climb of 0.3 percent for three-engine airplanes and 0.5 percent for four-engine airplanes." 14:14:1.0.1.3.13.2.75.20,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.125 Landing.,FAA,,,"[Amdt. 25-121, 72 FR 44666; Aug. 8, 2007; 72 FR 50467, Aug. 31, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) The horizontal distance necessary to land and to come to a complete stop (or to a speed of approximately 3 knots for water landings) from a point 50 feet above the landing surface must be determined (for standard temperatures, at each weight, altitude, and wind within the operational limits established by the applicant for the airplane): (1) In non-icing conditions; and (2) In icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if V REF for icing conditions exceeds V REF for non-icing conditions by more than 5 knots CAS at the maximum landing weight. (b) In determining the distance in paragraph (a) of this section: (1) The airplane must be in the landing configuration. (2) A stabilized approach, with a calibrated airspeed of not less than V REF , must be maintained down to the 50-foot height. (i) In non-icing conditions, V REF may not be less than: (A) 1.23 V SR 0; (B) V MCL established under § 25.149(f); and (C) A speed that provides the maneuvering capability specified in § 25.143(h). (ii) In icing conditions, V REF may not be less than: (A) The speed determined in paragraph (b)(2)(i) of this section; (B) 1.23 V SR0 with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if that speed exceeds V REF selected for non-icing conditions by more than 5 knots CAS; and (C) A speed that provides the maneuvering capability specified in § 25.143(h) with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g). (3) Changes in configuration, power or thrust, and speed, must be made in accordance with the established procedures for service operation. (4) The landing must be made without excessive vertical acceleration, tendency to bounce, nose over, ground loop, porpoise, or water loop. (5) The landings may not require exceptional piloting skill or alertness. (c) For landplanes and amphibians, the landing distance on land must be determined on a level, smooth, dry, hard-surfaced runway. In addition— (1) The pressures on the wheel braking systems may not exceed those specified by the brake manufacturer; (2) The brakes may not be used so as to cause excessive wear of brakes or tires; and (3) Means other than wheel brakes may be used if that means— (i) Is safe and reliable; (ii) Is used so that consistent results can be expected in service; and (iii) Is such that exceptional skill is not required to control the airplane. (d) For seaplanes and amphibians, the landing distance on water must be determined on smooth water. (e) For skiplanes, the landing distance on snow must be determined on smooth, dry, snow. (f) The landing distance data must include correction factors for not more than 50 percent of the nominal wind components along the landing path opposite to the direction of landing, and not less than 150 percent of the nominal wind components along the landing path in the direction of landing. (g) If any device is used that depends on the operation of any engine, and if the landing distance would be noticeably increased when a landing is made with that engine inoperative, the landing distance must be determined with that engine inoperative unless the use of compensating means will result in a landing distance not more than that with each engine operating." 14:14:1.0.1.3.13.2.75.8,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.101 General.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 25-92, 63 FR 8318, Feb. 18, 1998]","(a) Unless otherwise prescribed, airplanes must meet the applicable performance requirements of this subpart for ambient atmospheric conditions and still air. (b) The performance, as affected by engine power or thrust, must be based on the following relative humidities; (1) For turbine engine powered airplanes, a relative humidity of— (i) 80 percent, at and below standard temperatures; and (ii) 34 percent, at and above standard temperatures plus 50 °F. Between these two temperatures, the relative humidity must vary linearly. (2) For reciprocating engine powered airplanes, a relative humidity of 80 percent in a standard atmosphere. Engine power corrections for vapor pressure must be made in accordance with the following table: (c) The performance must correspond to the propulsive thrust available under the particular ambient atmospheric conditions, the particular flight condition, and the relative humidity specified in paragraph (b) of this section. The available propulsive thrust must correspond to engine power or thrust, not exceeding the approved power or thrust less— (1) Installation losses; and (2) The power or equivalent thrust absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and the particular flight condition. (d) Unless otherwise prescribed, the applicant must select the takeoff, en route, approach, and landing configurations for the airplane. (e) The airplane configurations may vary with weight, altitude, and temperature, to the extent they are compatible with the operating procedures required by paragraph (f) of this section. (f) Unless otherwise prescribed, in determining the accelerate-stop distances, takeoff flight paths, takeoff distances, and landing distances, changes in the airplane's configuration, speed, power, and thrust, must be made in accordance with procedures established by the applicant for operation in service. (g) Procedures for the execution of balked landings and missed approaches associated with the conditions prescribed in §§ 25.119 and 25.121(d) must be established. (h) The procedures established under paragraphs (f) and (g) of this section must— (1) Be able to be consistently executed in service by crews of average skill; (2) Use methods or devices that are safe and reliable; and (3) Include allowance for any time delays, in the execution of the procedures, that may reasonably be expected in service. (i) The accelerate-stop and landing distances prescribed in §§ 25.109 and 25.125, respectively, must be determined with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range." 14:14:1.0.1.3.13.2.75.9,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.103 Stall speed.,FAA,,,"[Doc. No. 28404, 67 FR 70825, Nov. 26, 2002, as amended by Amdt. 25-121, 72 FR 44665, Aug. 8, 2007]","(a) The reference stall speed, V SR , is a calibrated airspeed defined by the applicant. V SR may not be less than a 1-g stall speed. V SR is expressed as: where: V CL MAX = Calibrated airspeed obtained when the load factor-corrected lift coefficient is first a maximum during the maneuver prescribed in paragraph (c) of this section. In addition, when the maneuver is limited by a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher), V CL MAX may not be less than the speed existing at the instant the device operates; n ZW = Load factor normal to the flight path at V CL MAX W = Airplane gross weight; S = Aerodynamic reference wing area; and q = Dynamic pressure. where: V CL MAX = Calibrated airspeed obtained when the load factor-corrected lift coefficient is first a maximum during the maneuver prescribed in paragraph (c) of this section. In addition, when the maneuver is limited by a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher), V CL MAX may not be less than the speed existing at the instant the device operates; n ZW = Load factor normal to the flight path at V CL MAX W = Airplane gross weight; S = Aerodynamic reference wing area; and q = Dynamic pressure. (b) V CL MAX is determined with: (1) Engines idling, or, if that resultant thrust causes an appreciable decrease in stall speed, not more than zero thrust at the stall speed; (2) Propeller pitch controls (if applicable) in the takeoff position; (3) The airplane in other respects (such as flaps, landing gear, and ice accretions) in the condition existing in the test or performance standard in which V SR is being used; (4) The weight used when V SR is being used as a factor to determine compliance with a required performance standard; (5) The center of gravity position that results in the highest value of reference stall speed; and (6) The airplane trimmed for straight flight at a speed selected by the applicant, but not less than 1.13V SR and not greater than 1.3V SR . (c) Starting from the stabilized trim condition, apply the longitudinal control to decelerate the airplane so that the speed reduction does not exceed one knot per second. (d) In addition to the requirements of paragraph (a) of this section, when a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher) is installed, the reference stall speed, V SR , may not be less than 2 knots or 2 percent, whichever is greater, above the speed at which the device operates." 14:14:1.0.1.3.13.2.76.21,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.143 General.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 25-108, 67 FR 70826, Nov. 26, 2002; Amdt. 25-121, 72 FR 44667, Aug. 8, 2007; Amdt. 25-129, 74 FR 38339, Aug. 3, 2009; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) The airplane must be safely controllable and maneuverable during— (1) Takeoff; (2) Climb; (3) Level flight; (4) Descent; and (5) Landing. (b) It must be possible to make a smooth transition from one flight condition to any other flight condition without exceptional piloting skill, alertness, or strength, and without danger of exceeding the airplane limit-load factor under any probable operating conditions, including— (1) The sudden failure of the critical engine; (2) For airplanes with three or more engines, the sudden failure of the second critical engine when the airplane is in the en route, approach, or landing configuration and is trimmed with the critical engine inoperative; and (3) Configuration changes, including deployment or retraction of deceleration devices. (c) The airplane must be shown to be safely controllable and maneuverable with the most critical of the ice accretion(s) appropriate to the phase of flight as defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), and with the critical engine inoperative and its propeller (if applicable) in the minimum drag position: (1) At the minimum V 2 for takeoff; (2) During an approach and go-around; and (3) During an approach and landing. (d) The following table prescribes, for conventional wheel type controls, the maximum control forces permitted during the testing required by paragraph (a) through (c) of this section: (e) Approved operating procedures or conventional operating practices must be followed when demonstrating compliance with the control force limitations for short term application that are prescribed in paragraph (d) of this section. The airplane must be in trim, or as near to being in trim as practical, in the preceding steady flight condition. For the takeoff condition, the airplane must be trimmed according to the approved operating procedures. (f) When demonstrating compliance with the control force limitations for long term application that are prescribed in paragraph (d) of this section, the airplane must be in trim, or as near to being in trim as practical. (g) When maneuvering at a constant airspeed or Mach number (up to V FC /M FC ), the stick forces and the gradient of the stick force versus maneuvering load factor must lie within satisfactory limits. The stick forces must not be so great as to make excessive demands on the pilot's strength when maneuvering the airplane, and must not be so low that the airplane can easily be overstressed inadvertently. Changes of gradient that occur with changes of load factor must not cause undue difficulty in maintaining control of the airplane, and local gradients must not be so low as to result in a danger of overcontrolling. (h) The maneuvering capabilities in a constant speed coordinated turn at forward center of gravity, as specified in the following table, must be free of stall warning or other characteristics that might interfere with normal maneuvering: 1 A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the minimum climb gradient specified in § 25.121 for the flight condition. 2 Airspeed approved for all-engines-operating initial climb. 3 That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the remaining engines, would result in the thrust or power specified for the takeoff condition at V 2 , or any lesser thrust or power setting that is used for all-engines-operating initial climb procedures. (i) When demonstrating compliance with § 25.143 in icing conditions— (1) Controllability must be demonstrated with the most critical of the ice accretion(s) for the particular flight phase as defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g); (2) It must be shown that a push force is required throughout a pushover maneuver down to a zero g load factor, or the lowest load factor obtainable if limited by elevator power or other design characteristic of the flight control system. It must be possible to promptly recover from the maneuver without exceeding a pull control force of 50 pounds; and (3) Any changes in force that the pilot must apply to the pitch control to maintain speed with increasing sideslip angle must be steadily increasing with no force reversals, unless the change in control force is gradual and easily controllable by the pilot without using exceptional piloting skill, alertness, or strength. (j) For flight in icing conditions before the ice protection system has been activated and is performing its intended function, it must be demonstrated in flight with the most critical of the ice accretion(s) defined in Appendix C, part II, paragraph (e) of this part and Appendix O, part II, paragraph (d) of this part, as applicable, in accordance with § 25.21(g), that: (1) The airplane is controllable in a pull-up maneuver up to 1.5 g load factor; and (2) There is no pitch control force reversal during a pushover maneuver down to 0.5 g load factor." 14:14:1.0.1.3.13.2.76.22,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.145 Longitudinal control.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29774, July 20, 1990; Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 25-98, 64 FR 6164, Feb. 8, 1999; 64 FR 10740, Mar. 5, 1999; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]","(a) It must be possible, at any point between the trim speed prescribed in § 25.103(b)(6) and stall identification (as defined in § 25.201(d)), to pitch the nose downward so that the acceleration to this selected trim speed is prompt with (1) The airplane trimmed at the trim speed prescribed in § 25.103(b)(6); (2) The landing gear extended; (3) The wing flaps (i) retracted and (ii) extended; and (4) Power (i) off and (ii) at maximum continuous power on the engines. (b) With the landing gear extended, no change in trim control, or exertion of more than 50 pounds control force (representative of the maximum short term force that can be applied readily by one hand) may be required for the following maneuvers: (1) With power off, flaps retracted, and the airplane trimmed at 1.3 V SR1 , extend the flaps as rapidly as possible while maintaining the airspeed at approximately 30 percent above the reference stall speed existing at each instant throughout the maneuver. (2) Repeat paragraph (b)(1) except initially extend the flaps and then retract them as rapidly as possible. (3) Repeat paragraph (b)(2), except at the go-around power or thrust setting. (4) With power off, flaps retracted, and the airplane trimmed at 1.3 V SR1 , rapidly set go-around power or thrust while maintaining the same airspeed. (5) Repeat paragraph (b)(4) except with flaps extended. (6) With power off, flaps extended, and the airplane trimmed at 1.3 V SR1 , obtain and maintain airspeeds between V SW and either 1.6 V SR1 or V FE , whichever is lower. (c) It must be possible, without exceptional piloting skill, to prevent loss of altitude when complete retraction of the high lift devices from any position is begun during steady, straight, level flight at 1.08 V SR1 for propeller powered airplanes, or 1.13 V SR1 for turbojet powered airplanes, with— (1) Simultaneous movement of the power or thrust controls to the go-around power or thrust setting; (2) The landing gear extended; and (3) The critical combinations of landing weights and altitudes. (d) If gated high-lift device control positions are provided, paragraph (c) of this section applies to retractions of the high-lift devices from any position from the maximum landing position to the first gated position, between gated positions, and from the last gated position to the fully retracted position. The requirements of paragraph (c) of this section also apply to retractions from each approved landing position to the control position(s) associated with the high-lift device configuration(s) used to establish the go-around procedure(s) from that landing position. In addition, the first gated control position from the maximum landing position must correspond with a configuration of the high-lift devices used to establish a go-around procedure from a landing configuration. Each gated control position must require a separate and distinct motion of the control to pass through the gated position and must have features to prevent inadvertent movement of the control through the gated position. It must only be possible to make this separate and distinct motion once the control has reached the gated position." 14:14:1.0.1.3.13.2.76.23,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.147 Directional and lateral control.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 2004]","(a) Directional control; general. It must be possible, with the wings level, to yaw into the operative engine and to safely make a reasonably sudden change in heading of up to 15 degrees in the direction of the critical inoperative engine. This must be shown at 1.3 V S R1 for heading changes up to 15 degrees (except that the heading change at which the rudder pedal force is 150 pounds need not be exceeded), and with— (1) The critical engine inoperative and its propeller in the minimum drag position; (2) The power required for level flight at 1.3 V S R1, but not more than maximum continuous power; (3) The most unfavorable center of gravity; (4) Landing gear retracted; (5) Flaps in the approach position; and (6) Maximum landing weight. (b) Directional control; airplanes with four or more engines. Airplanes with four or more engines must meet the requirements of paragraph (a) of this section except that— (1) The two critical engines must be inoperative with their propellers (if applicable) in the minimum drag position; (2) [Reserved] (3) The flaps must be in the most favorable climb position. (c) Lateral control; general. It must be possible to make 20° banked turns, with and against the inoperative engine, from steady flight at a speed equal to 1.3 V S R1, with— (1) The critical engine inoperative and its propeller (if applicable) in the minimum drag position; (2) The remaining engines at maximum continuous power; (3) The most unfavorable center of gravity; (4) Landing gear (i) retracted and (ii) extended; (5) Flaps in the most favorable climb position; and (6) Maximum takeoff weight. (d) Lateral control; roll capability. With the critical engine inoperative, roll response must allow normal maneuvers. Lateral control must be sufficient, at the speeds likely to be used with one engine inoperative, to provide a roll rate necessary for safety without excessive control forces or travel. (e) Lateral control; airplanes with four or more engines. Airplanes with four or more engines must be able to make 20° banked turns, with and against the inoperative engines, from steady flight at a speed equal to 1.3 V S R1, with maximum continuous power, and with the airplane in the configuration prescribed by paragraph (b) of this section. (f) Lateral control; all engines operating. With the engines operating, roll response must allow normal maneuvers (such as recovery from upsets produced by gusts and the initiation of evasive maneuvers). There must be enough excess lateral control in sideslips (up to sideslip angles that might be required in normal operation), to allow a limited amount of maneuvering and to correct for gusts. Lateral control must be enough at any speed up to V FC / M FC to provide a peak roll rate necessary for safety, without excessive control forces or travel." 14:14:1.0.1.3.13.2.76.24,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.149 Minimum control speed.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2321, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25-84, 60 FR 30749, June 9, 1995; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]","(a) In establishing the minimum control speeds required by this section, the method used to simulate critical engine failure must represent the most critical mode of powerplant failure with respect to controllability expected in service. (b) V MC is the calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with that engine still inoperative and maintain straight flight with an angle of bank of not more than 5 degrees. (c) V MC may not exceed 1.13 V SR with— (1) Maximum available takeoff power or thrust on the engines; (2) The most unfavorable center of gravity; (3) The airplane trimmed for takeoff; (4) The maximum sea level takeoff weight (or any lesser weight necessary to show V MC ); (5) The airplane in the most critical takeoff configuration existing along the flight path after the airplane becomes airborne, except with the landing gear retracted; (6) The airplane airborne and the ground effect negligible; and (7) If applicable, the propeller of the inoperative engine— (i) Windmilling; (ii) In the most probable position for the specific design of the propeller control; or (iii) Feathered, if the airplane has an automatic feathering device acceptable for showing compliance with the climb requirements of § 25.121. (d) The rudder forces required to maintain control at V MC may not exceed 150 pounds nor may it be necessary to reduce power or thrust of the operative engines. During recovery, the airplane may not assume any dangerous attitude or require exceptional piloting skill, alertness, or strength to prevent a heading change of more than 20 degrees. (e) V MCG , the minimum control speed on the ground, is the calibrated airspeed during the takeoff run at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane using the rudder control alone (without the use of nosewheel steering), as limited by 150 pounds of force, and the lateral control to the extent of keeping the wings level to enable the takeoff to be safely continued using normal piloting skill. In the determination of V MCG , assuming that the path of the airplane accelerating with all engines operating is along the centerline of the runway, its path from the point at which the critical engine is made inoperative to the point at which recovery to a direction parallel to the centerline is completed may not deviate more than 30 feet laterally from the centerline at any point. V MCG must be established with— (1) The airplane in each takeoff configuration or, at the option of the applicant, in the most critical takeoff configuration; (2) Maximum available takeoff power or thrust on the operating engines; (3) The most unfavorable center of gravity; (4) The airplane trimmed for takeoff; and (5) The most unfavorable weight in the range of takeoff weights. (f) V MCL , the minimum control speed during approach and landing with all engines operating, is the calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with that engine still inoperative, and maintain straight flight with an angle of bank of not more than 5 degrees. V MCL must be established with— (1) The airplane in the most critical configuration (or, at the option of the applicant, each configuration) for approach and landing with all engines operating; (2) The most unfavorable center of gravity; (3) The airplane trimmed for approach with all engines operating; (4) The most favorable weight, or, at the option of the applicant, as a function of weight; (5) For propeller airplanes, the propeller of the inoperative engine in the position it achieves without pilot action, assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle; and (6) Go-around power or thrust setting on the operating engine(s). (g) For airplanes with three or more engines, V MCL-2 , the minimum control speed during approach and landing with one critical engine inoperative, is the calibrated airspeed at which, when a second critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with both engines still inoperative, and maintain straight flight with an angle of bank of not more than 5 degrees. V MCL-2 must be established with— (1) The airplane in the most critical configuration (or, at the option of the applicant, each configuration) for approach and landing with one critical engine inoperative; (2) The most unfavorable center of gravity; (3) The airplane trimmed for approach with one critical engine inoperative; (4) The most unfavorable weight, or, at the option of the applicant, as a function of weight; (5) For propeller airplanes, the propeller of the more critical inoperative engine in the position it achieves without pilot action, assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle, and the propeller of the other inoperative engine feathered; (6) The power or thrust on the operating engine(s) necessary to maintain an approach path angle of three degrees when one critical engine is inoperative; and (7) The power or thrust on the operating engine(s) rapidly changed, immediately after the second critical engine is made inoperative, from the power or thrust prescribed in paragraph (g)(6) of this section to— (i) Minimum power or thrust; and (ii) Go-around power or thrust setting. (h) In demonstrations of V MCL and V MCL-2 — (1) The rudder force may not exceed 150 pounds; (2) The airplane may not exhibit hazardous flight characteristics or require exceptional piloting skill, alertness, or strength; (3) Lateral control must be sufficient to roll the airplane, from an initial condition of steady flight, through an angle of 20 degrees in the direction necessary to initiate a turn away from the inoperative engine(s), in not more than 5 seconds; and (4) For propeller airplanes, hazardous flight characteristics must not be exhibited due to any propeller position achieved when the engine fails or during any likely subsequent movements of the engine or propeller controls." 14:14:1.0.1.3.13.2.77.25,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.161 Trim.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 2004]","(a) General. Each airplane must meet the trim requirements of this section after being trimmed, and without further pressure upon, or movement of, either the primary controls or their corresponding trim controls by the pilot or the automatic pilot. (b) Lateral and directional trim. The airplane must maintain lateral and directional trim with the most adverse lateral displacement of the center of gravity within the relevant operating limitations, during normally expected conditions of operation (including operation at any speed from 1.3 V SR 1 to V MO /M MO ). (c) Longitudinal trim. The airplane must maintain longitudinal trim during— (1) A climb with maximum continuous power at a speed not more than 1.3 V SR 1 , with the landing gear retracted, and the flaps (i) retracted and (ii) in the takeoff position; (2) Either a glide with power off at a speed not more than 1.3 V SR1 , or an approach within the normal range of approach speeds appropriate to the weight and configuration with power settings corresponding to a 3 degree glidepath, whichever is the most severe, with the landing gear extended, the wing flaps (i) retracted and (ii) extended, and with the most unfavorable combination of center of gravity position and weight approved for landing; and (3) Level flight at any speed from 1.3 V SR 1 , to V MO /M MO, with the landing gear and flaps retracted, and from 1.3 V SR 1 to V LE with the landing gear extended. (d) Longitudinal, directional, and lateral trim. The airplane must maintain longitudinal, directional, and lateral trim (and for the lateral trim, the angle of bank may not exceed five degrees) at 1.3 V SR 1 during climbing flight with— (1) The critical engine inoperative; (2) The remaining engines at maximum continuous power; and (3) The landing gear and flaps retracted. (e) Airplanes with four or more engines. Each airplane with four or more engines must also maintain trim in rectilinear flight with the most unfavorable center of gravity and at the climb speed, configuration, and power required by § 25.123(a) for the purpose of establishing the en route flight paths with two engines inoperative." 14:14:1.0.1.3.13.2.78.26,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.171 General.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR 13117, Oct. 15, 1965]","The airplane must be longitudinally, directionally, and laterally stable in accordance with the provisions of §§ 25.173 through 25.177. In addition, suitable stability and control feel (static stability) is required in any condition normally encountered in service, if flight tests show it is necessary for safe operation." 14:14:1.0.1.3.13.2.78.27,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.173 Static longitudinal stability.,FAA,,,"[Amdt. 25-7, 30 FR 13117, Oct. 15, 1965]","Under the conditions specified in § 25.175, the characteristics of the elevator control forces (including friction) must be as follows: (a) A pull must be required to obtain and maintain speeds below the specified trim speed, and a push must be required to obtain and maintain speeds above the specified trim speed. This must be shown at any speed that can be obtained except speeds higher than the landing gear or wing flap operating limit speeds or V FC /M FC, whichever is appropriate, or lower than the minimum speed for steady unstalled flight. (b) The airspeed must return to within 10 percent of the original trim speed for the climb, approach, and landing conditions specified in § 25.175 (a), (c), and (d), and must return to within 7.5 percent of the original trim speed for the cruising condition specified in § 25.175(b), when the control force is slowly released from any speed within the range specified in paragraph (a) of this section. (c) The average gradient of the stable slope of the stick force versus speed curve may not be less than 1 pound for each 6 knots. (d) Within the free return speed range specified in paragraph (b) of this section, it is permissible for the airplane, without control forces, to stabilize on speeds above or below the desired trim speeds if exceptional attention on the part of the pilot is not required to return to and maintain the desired trim speed and altitude." 14:14:1.0.1.3.13.2.78.28,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.175 Demonstration of static longitudinal stability.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR 13117, Oct. 15, 1965; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-115, 69 FR 40527, July 2, 2004]","Static longitudinal stability must be shown as follows: (a) Climb. The stick force curve must have a stable slope at speeds between 85 and 115 percent of the speed at which the airplane— (1) Is trimmed, with— (i) Wing flaps retracted; (ii) Landing gear retracted; (iii) Maximum takeoff weight; and (iv) 75 percent of maximum continuous power for reciprocating engines or the maximum power or thrust selected by the applicant as an operating limitation for use during climb for turbine engines; and (2) Is trimmed at the speed for best rate-of-climb except that the speed need not be less than 1.3 V SR 1 . (b) Cruise. Static longitudinal stability must be shown in the cruise condition as follows: (1) With the landing gear retracted at high speed, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than V FC /M FC, nor speeds that require a stick force of more than 50 pounds), with— (i) The wing flaps retracted; (ii) The center of gravity in the most adverse position (see § 25.27); (iii) The most critical weight between the maximum takeoff and maximum landing weights; (iv) 75 percent of maximum continuous power for reciprocating engines or for turbine engines, the maximum cruising power selected by the applicant as an operating limitation (see § 25.1521), except that the power need not exceed that required at V MO / M MO ; and (v) The airplane trimmed for level flight with the power required in paragraph (b)(1)(iv) of this section. (2) With the landing gear retracted at low speed, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than the minimum speed of the applicable speed range prescribed in paragraph (b)(1), nor speeds that require a stick force of more than 50 pounds), with— (i) Wing flaps, center of gravity position, and weight as specified in paragraph (b)(1) of this section; (ii) Power required for level flight at a speed equal to ( V MO + 1.3 V SR 1 )/2; and (iii) The airplane trimmed for level flight with the power required in paragraph (b)(2)(ii) of this section. (3) With the landing gear extended, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than V LE, nor speeds that require a stick force of more than 50 pounds), with— (i) Wing flap, center of gravity position, and weight as specified in paragraph (b)(1) of this section; (ii) 75 percent of maximum continuous power for reciprocating engines or, for turbine engines, the maximum cruising power selected by the applicant as an operating limitation, except that the power need not exceed that required for level flight at V LE ; and (iii) The aircraft trimmed for level flight with the power required in paragraph (b)(3)(ii) of this section. (c) Approach. The stick force curve must have a stable slope at speeds between V SW and 1.7 V SR 1 , with— (1) Wing flaps in the approach position; (2) Landing gear retracted; (3) Maximum landing weight; and (4) The airplane trimmed at 1.3 V SR 1 with enough power to maintain level flight at this speed. (d) Landing. The stick force curve must have a stable slope, and the stick force may not exceed 80 pounds, at speeds between V SW and 1.7 V SR 0 with— (1) Wing flaps in the landing position; (2) Landing gear extended; (3) Maximum landing weight; (4) The airplane trimmed at 1.3 V SR0 with— (i) Power or thrust off, and (ii) Power or thrust for level flight. (5) The airplane trimmed at 1.3 V SR 0 with power or thrust off." 14:14:1.0.1.3.13.2.78.29,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.177 Static lateral-directional stability.,FAA,,,"[Amdt. 25-135, 76 FR 74654, Dec. 1, 2011]","(a) The static directional stability (as shown by the tendency to recover from a skid with the rudder free) must be positive for any landing gear and flap position and symmetric power condition, at speeds from 1.13 V SR1 , up to V FE , V LE , or V FC /M FC (as appropriate for the airplane configuration). (b) The static lateral stability (as shown by the tendency to raise the low wing in a sideslip with the aileron controls free) for any landing gear and flap position and symmetric power condition, may not be negative at any airspeed (except that speeds higher than V FE need not be considered for flaps extended configurations nor speeds higher than V LE for landing gear extended configurations) in the following airspeed ranges: (1) From 1.13 V SR1 to V MO /M MO . (2) From V MO /M MO to V FC /M FC , unless the divergence is— (i) Gradual; (ii) Easily recognizable by the pilot; and (iii) Easily controllable by the pilot. (c) The following requirement must be met for the configurations and speed specified in paragraph (a) of this section. In straight, steady sideslips over the range of sideslip angles appropriate to the operation of the airplane, the aileron and rudder control movements and forces must be substantially proportional to the angle of sideslip in a stable sense. This factor of proportionality must lie between limits found necessary for safe operation. The range of sideslip angles evaluated must include those sideslip angles resulting from the lesser of: (1) One-half of the available rudder control input; and (2) A rudder control force of 180 pounds. (d) For sideslip angles greater than those prescribed by paragraph (c) of this section, up to the angle at which full rudder control is used or a rudder control force of 180 pounds is obtained, the rudder control forces may not reverse, and increased rudder deflection must be needed for increased angles of sideslip. Compliance with this requirement must be shown using straight, steady sideslips, unless full lateral control input is achieved before reaching either full rudder control input or a rudder control force of 180 pounds; a straight, steady sideslip need not be maintained after achieving full lateral control input. This requirement must be met at all approved landing gear and flap positions for the range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating." 14:14:1.0.1.3.13.2.78.30,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.181 Dynamic stability.,FAA,,,"[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25-72, 55 FR 29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]","(a) Any short period oscillation, not including combined lateral-directional oscillations, occurring between 1.13 V SR and maximum allowable speed appropriate to the configuration of the airplane must be heavily damped with the primary controls— (1) Free; and (2) In a fixed position. (b) Any combined lateral-directional oscillations (“Dutch roll”) occurring between 1.13 V SR and maximum allowable speed appropriate to the configuration of the airplane must be positively damped with controls free, and must be controllable with normal use of the primary controls without requiring exceptional pilot skill." 14:14:1.0.1.3.13.2.79.31,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.201 Stall demonstration.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 60 FR 30750, June 9, 1995; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002]","(a) Stalls must be shown in straight flight and in 30 degree banked turns with— (1) Power off; and (2) The power necessary to maintain level flight at 1.5 V SR1 (where V SR1 corresponds to the reference stall speed at maximum landing weight with flaps in the approach position and the landing gear retracted). (b) In each condition required by paragraph (a) of this section, it must be possible to meet the applicable requirements of § 25.203 with— (1) Flaps, landing gear, and deceleration devices in any likely combination of positions approved for operation; (2) Representative weights within the range for which certification is requested; (3) The most adverse center of gravity for recovery; and (4) The airplane trimmed for straight flight at the speed prescribed in § 25.103(b)(6). (c) The following procedures must be used to show compliance with § 25.203; (1) Starting at a speed sufficiently above the stalling speed to ensure that a steady rate of speed reduction can be established, apply the longitudinal control so that the speed reduction does not exceed one knot per second until the airplane is stalled. (2) In addition, for turning flight stalls, apply the longitudinal control to achieve airspeed deceleration rates up to 3 knots per second. (3) As soon as the airplane is stalled, recover by normal recovery techniques. (d) The airplane is considered stalled when the behavior of the airplane gives the pilot a clear and distinctive indication of an acceptable nature that the airplane is stalled. Acceptable indications of a stall, occurring either individually or in combination, are— (1) A nose-down pitch that cannot be readily arrested; (2) Buffeting, of a magnitude and severity that is a strong and effective deterrent to further speed reduction; or (3) The pitch control reaches the aft stop and no further increase in pitch attitude occurs when the control is held full aft for a short time before recovery is initiated." 14:14:1.0.1.3.13.2.79.32,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.203 Stall characteristics.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-84, 60 FR 30750, June 9, 1995]","(a) It must be possible to produce and to correct roll and yaw by unreversed use of the aileron and rudder controls, up to the time the airplane is stalled. No abnormal nose-up pitching may occur. The longitudinal control force must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and to recover from a stall by normal use of the controls. (b) For level wing stalls, the roll occurring between the stall and the completion of the recovery may not exceed approximately 20 degrees. (c) For turning flight stalls, the action of the airplane after the stall may not be so violent or extreme as to make it difficult, with normal piloting skill, to effect a prompt recovery and to regain control of the airplane. The maximum bank angle that occurs during the recovery may not exceed— (1) Approximately 60 degrees in the original direction of the turn, or 30 degrees in the opposite direction, for deceleration rates up to 1 knot per second; and (2) Approximately 90 degrees in the original direction of the turn, or 60 degrees in the opposite direction, for deceleration rates in excess of 1 knot per second." 14:14:1.0.1.3.13.2.79.33,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.207 Stall warning.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR 13118, Oct. 15, 1965; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-121, 72 FR 44668, Aug. 8, 2007; Amdt. 25-129, 74 FR 38339, Aug. 3, 2009; Amdt. 25-140, 79 FR 65526, Nov. 4, 2014]","(a) Stall warning with sufficient margin to prevent inadvertent stalling with the flaps and landing gear in any normal position must be clear and distinctive to the pilot in straight and turning flight. (b) The warning must be furnished either through the inherent aerodynamic qualities of the airplane or by a device that will give clearly distinguishable indications under expected conditions of flight. However, a visual stall warning device that requires the attention of the crew within the cockpit is not acceptable by itself. If a warning device is used, it must provide a warning in each of the airplane configurations prescribed in paragraph (a) of this section at the speed prescribed in paragraphs (c) and (d) of this section. Except for the stall warning prescribed in paragraph (h)(3)(ii) of this section, the stall warning for flight in icing conditions must be provided by the same means as the stall warning for flight in non-icing conditions. (c) When the speed is reduced at rates not exceeding one knot per second, stall warning must begin, in each normal configuration, at a speed, V SW , exceeding the speed at which the stall is identified in accordance with § 25.201(d) by not less than five knots or five percent CAS, whichever is greater. Once initiated, stall warning must continue until the angle of attack is reduced to approximately that at which stall warning began. (d) In addition to the requirement of paragraph (c) of this section, when the speed is reduced at rates not exceeding one knot per second, in straight flight with engines idling and at the center-of-gravity position specified in § 25.103(b)(5), V SW , in each normal configuration, must exceed V SR by not less than three knots or three percent CAS, whichever is greater. (e) In icing conditions, the stall warning margin in straight and turning flight must be sufficient to allow the pilot to prevent stalling (as defined in § 25.201(d)) when the pilot starts a recovery maneuver not less than three seconds after the onset of stall warning. When demonstrating compliance with this paragraph, the pilot must perform the recovery maneuver in the same way as for the airplane in non-icing conditions. Compliance with this requirement must be demonstrated in flight with the speed reduced at rates not exceeding one knot per second, with— (1) The most critical of the takeoff ice and final takeoff ice accretions defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for each configuration used in the takeoff phase of flight; (2) The most critical of the en route ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the en route configuration; (3) The most critical of the holding ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the holding configuration(s); (4) The most critical of the approach ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the approach configuration(s); and (5) The most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the landing and go-around configuration(s). (f) The stall warning margin must be sufficient in both non-icing and icing conditions to allow the pilot to prevent stalling when the pilot starts a recovery maneuver not less than one second after the onset of stall warning in slow-down turns with at least 1.5 g load factor normal to the flight path and airspeed deceleration rates of at least 2 knots per second. When demonstrating compliance with this paragraph for icing conditions, the pilot must perform the recovery maneuver in the same way as for the airplane in non-icing conditions. Compliance with this requirement must be demonstrated in flight with— (1) The flaps and landing gear in any normal position; (2) The airplane trimmed for straight flight at a speed of 1.3 V SR ; and (3) The power or thrust necessary to maintain level flight at 1.3 V SR . (g) Stall warning must also be provided in each abnormal configuration of the high lift devices that is likely to be used in flight following system failures (including all configurations covered by Airplane Flight Manual procedures). (h) The following stall warning margin is required for flight in icing conditions before the ice protection system has been activated and is performing its intended function. Compliance must be shown using the most critical of the ice accretion(s) defined in Appendix C, part II, paragraph (e) of this part and Appendix O, part II, paragraph (d) of this part, as applicable, in accordance with § 25.21(g). The stall warning margin in straight and turning flight must be sufficient to allow the pilot to prevent stalling without encountering any adverse flight characteristics when: (1) The speed is reduced at rates not exceeding one knot per second; (2) The pilot performs the recovery maneuver in the same way as for flight in non-icing conditions; and (3) The recovery maneuver is started no earlier than: (i) One second after the onset of stall warning if stall warning is provided by the same means as for flight in non-icing conditions; or (ii) Three seconds after the onset of stall warning if stall warning is provided by a different means than for flight in non-icing conditions. (i) In showing compliance with paragraph (h) of this section, if stall warning is provided by a different means in icing conditions than for non-icing conditions, compliance with § 25.203 must be shown using the accretion defined in appendix C, part II(e) of this part. Compliance with this requirement must be shown using the demonstration prescribed by § 25.201, except that the deceleration rates of § 25.201(c)(2) need not be demonstrated." 14:14:1.0.1.3.13.2.80.34,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.231 Longitudinal stability and control.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-108, 67 FR 70828, Nov. 26, 2002]","(a) Landplanes may have no uncontrollable tendency to nose over in any reasonably expected operating condition or when rebound occurs during landing or takeoff. In addition— (1) Wheel brakes must operate smoothly and may not cause any undue tendency to nose over; and (2) If a tail-wheel landing gear is used, it must be possible, during the takeoff ground run on concrete, to maintain any attitude up to thrust line level, at 75 percent of V SR 1 . (b) For seaplanes and amphibians, the most adverse water conditions safe for takeoff, taxiing, and landing, must be established." 14:14:1.0.1.3.13.2.80.35,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.233 Directional stability and control.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998; Amdt. 25-108, 67 FR 70828, Nov. 26, 2002]","(a) There may be no uncontrollable ground-looping tendency in 90° cross winds, up to a wind velocity of 20 knots or 0.2 V SR 0 , whichever is greater, except that the wind velocity need not exceed 25 knots at any speed at which the airplane may be expected to be operated on the ground. This may be shown while establishing the 90° cross component of wind velocity required by § 25.237. (b) Landplanes must be satisfactorily controllable, without exceptional piloting skill or alertness, in power-off landings at normal landing speed, without using brakes or engine power to maintain a straight path. This may be shown during power-off landings made in conjunction with other tests. (c) The airplane must have adequate directional control during taxiing. This may be shown during taxiing prior to takeoffs made in conjunction with other tests." 14:14:1.0.1.3.13.2.80.36,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.235 Taxiing condition.,FAA,,,,The shock absorbing mechanism may not damage the structure of the airplane when the airplane is taxied on the roughest ground that may reasonably be expected in normal operation. 14:14:1.0.1.3.13.2.80.37,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.237 Wind velocities.,FAA,,,"[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-121, 72 FR 44668, Aug. 8, 2007; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014]","(a) For land planes and amphibians, the following applies: (1) A 90-degree cross component of wind velocity, demonstrated to be safe for takeoff and landing, must be established for dry runways and must be at least 20 knots or 0.2 V SR0 , whichever is greater, except that it need not exceed 25 knots. (2) The crosswind component for takeoff established without ice accretions is valid in icing conditions. (3) The landing crosswind component must be established for: (i) Non-icing conditions, and (ii) Icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g). (b) For seaplanes and amphibians, the following applies: (1) A 90-degree cross component of wind velocity, up to which takeoff and landing is safe under all water conditions that may reasonably be expected in normal operation, must be established and must be at least 20 knots or 0.2 V SR 0 , whichever is greater, except that it need not exceed 25 knots. (2) A wind velocity, for which taxiing is safe in any direction under all water conditions that may reasonably be expected in normal operation, must be established and must be at least 20 knots or 0.2 V SR0 , whichever is greater, except that it need not exceed 25 knots." 14:14:1.0.1.3.13.2.80.38,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,"§ 25.239 Spray characteristics, control, and stability on water.",FAA,,,,"(a) For seaplanes and amphibians, during takeoff, taxiing, and landing, and in the conditions set forth in paragraph (b) of this section, there may be no— (1) Spray characteristics that would impair the pilot's view, cause damage, or result in the taking in of an undue quantity of water; (2) Dangerously uncontrollable porpoising, bounding, or swinging tendency; or (3) Immersion of auxiliary floats or sponsons, wing tips, propeller blades, or other parts not designed to withstand the resulting water loads. (b) Compliance with the requirements of paragraph (a) of this section must be shown— (1) In water conditions, from smooth to the most adverse condition established in accordance with § 25.231; (2) In wind and cross-wind velocities, water currents, and associated waves and swells that may reasonably be expected in operation on water; (3) At speeds that may reasonably be expected in operation on water; (4) With sudden failure of the critical engine at any time while on water; and (5) At each weight and center of gravity position, relevant to each operating condition, within the range of loading conditions for which certification is requested. (c) In the water conditions of paragraph (b) of this section, and in the corresponding wind conditions, the seaplane or amphibian must be able to drift for five minutes with engines inoperative, aided, if necessary, by a sea anchor." 14:14:1.0.1.3.13.2.81.39,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.251 Vibration and buffeting.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 25-77, 57 FR 28949, June 29, 1992]","(a) The airplane must be demonstrated in flight to be free from any vibration and buffeting that would prevent continued safe flight in any likely operating condition. (b) Each part of the airplane must be demonstrated in flight to be free from excessive vibration under any appropriate speed and power conditions up to V DF /M DF . The maximum speeds shown must be used in establishing the operating limitations of the airplane in accordance with § 25.1505. (c) Except as provided in paragraph (d) of this section, there may be no buffeting condition, in normal flight, including configuration changes during cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the crew, or to cause structural damage. Stall warning buffeting within these limits is allowable. (d) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any speed up to V MO / M MO, except that stall warning buffeting is allowable. (e) For an airplane with M D greater than .6 or with a maximum operating altitude greater than 25,000 feet, the positive maneuvering load factors at which the onset of perceptible buffeting occurs must be determined with the airplane in the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the airplane is to be certificated. The envelopes of load factor, speed, altitude, and weight must provide a sufficient range of speeds and load factors for normal operations. Probable inadvertent excursions beyond the boundaries of the buffet onset envelopes may not result in unsafe conditions." 14:14:1.0.1.3.13.2.81.40,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.253 High-speed characteristics.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5671, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 25-84, 60 FR 30750, June 9, 1995; Amdt. 25-121, 72 FR 44668, Aug. 8, 2007; Amdt. 25-135, 76 FR 74654, Dec. 1, 2011; Amdt. 25-140,79 FR 65525, Nov. 4, 2014]","(a) Speed increase and recovery characteristics. The following speed increase and recovery characteristics must be met: (1) Operating conditions and characteristics likely to cause inadvertent speed increases (including upsets in pitch and roll) must be simulated with the airplane trimmed at any likely cruise speed up to V MO / M MO . These conditions and characteristics include gust upsets, inadvertent control movements, low stick force gradient in relation to control friction, passenger movement, leveling off from climb, and descent from Mach to airspeed limit altitudes. (2) Allowing for pilot reaction time after effective inherent or artificial speed warning occurs, it must be shown that the airplane can be recovered to a normal attitude and its speed reduced to V MO / M MO, without— (i) Exceptional piloting strength or skill; (ii) Exceeding V D / M D, V DF / M DF, or the structural limitations; and (iii) Buffeting that would impair the pilot's ability to read the instruments or control the airplane for recovery. (3) With the airplane trimmed at any speed up to V MO /M MO , there must be no reversal of the response to control input about any axis at any speed up to V DF /M DF . Any tendency to pitch, roll, or yaw must be mild and readily controllable, using normal piloting techniques. When the airplane is trimmed at V MO /M MO , the slope of the elevator control force versus speed curve need not be stable at speeds greater than V FC /M FC , but there must be a push force at all speeds up to V DF /M DF and there must be no sudden or excessive reduction of elevator control force as V DF /M DF is reached. (4) Adequate roll capability to assure a prompt recovery from a lateral upset condition must be available at any speed up to V DF /M DF . (5) With the airplane trimmed at V MO /M MO , extension of the speedbrakes over the available range of movements of the pilot's control, at all speeds above V MO /M MO , but not so high that V DF /M DF would be exceeded during the maneuver, must not result in: (i) An excessive positive load factor when the pilot does not take action to counteract the effects of extension; (ii) Buffeting that would impair the pilot's ability to read the instruments or control the airplane for recovery; or (iii) A nose down pitching moment, unless it is small. (b) Maximum speed for stability characteristics, V FC /M FC . V FC /M FC is the maximum speed at which the requirements of §§ 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), and 25.181 must be met with flaps and landing gear retracted. Except as noted in § 25.253(c), V FC /M FC may not be less than a speed midway between V MO /M MO and V DF /M DF , except that, for altitudes where Mach number is the limiting factor, M FC need not exceed the Mach number at which effective speed warning occurs. (c) Maximum speed for stability characteristics in icing conditions. The maximum speed for stability characteristics with the most critical of the ice accretions defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), at which the requirements of §§ 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), and 25.181 must be met, is the lower of: (1) 300 knots CAS; (2) V FC ; or (3) A speed at which it is demonstrated that the airframe will be free of ice accretion due to the effects of increased dynamic pressure." 14:14:1.0.1.3.13.2.81.41,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,B,Subpart B—Flight,,§ 25.255 Out-of-trim characteristics.,FAA,,,"[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]","(a) From an initial condition with the airplane trimmed at cruise speeds up to V MO /M MO, the airplane must have satisfactory maneuvering stability and controllability with the degree of out-of-trim in both the airplane nose-up and nose-down directions, which results from the greater of— (1) A three-second movement of the longitudinal trim system at its normal rate for the particular flight condition with no aerodynamic load (or an equivalent degree of trim for airplanes that do not have a power-operated trim system), except as limited by stops in the trim system, including those required by § 25.655(b) for adjustable stabilizers; or (2) The maximum mistrim that can be sustained by the autopilot while maintaining level flight in the high speed cruising condition. (b) In the out-of-trim condition specified in paragraph (a) of this section, when the normal acceleration is varied from + 1 g to the positive and negative values specified in paragraph (c) of this section— (1) The stick force vs. g curve must have a positive slope at any speed up to and including V FC /M FC ; and (2) At speeds between V FC /M FC and V DF /M DF the direction of the primary longitudinal control force may not reverse. (c) Except as provided in paragraphs (d) and (e) of this section, compliance with the provisions of paragraph (a) of this section must be demonstrated in flight over the acceleration range— (1) −1 g to + 2.5 g; or (2) 0 g to 2.0 g, and extrapolating by an acceptable method to −1 g and + 2.5 g. (d) If the procedure set forth in paragraph (c)(2) of this section is used to demonstrate compliance and marginal conditions exist during flight test with regard to reversal of primary longitudinal control force, flight tests must be accomplished from the normal acceleration at which a marginal condition is found to exist to the applicable limit specified in paragraph (b)(1) of this section. (e) During flight tests required by paragraph (a) of this section, the limit maneuvering load factors prescribed in §§ 25.333(b) and 25.337, and the maneuvering load factors associated with probable inadvertent excursions beyond the boundaries of the buffet onset envelopes determined under § 25.251(e), need not be exceeded. In addition, the entry speeds for flight test demonstrations at normal acceleration values less than 1 g must be limited to the extent necessary to accomplish a recovery without exceeding V DF /M DF . (f) In the out-of-trim condition specified in paragraph (a) of this section, it must be possible from an overspeed condition at V DF /M DF to produce at least 1.5 g for recovery by applying not more than 125 pounds of longitudinal control force using either the primary longitudinal control alone or the primary longitudinal control and the longitudinal trim system. If the longitudinal trim is used to assist in producing the required load factor, it must be shown at V DF /M DF that the longitudinal trim can be actuated in the airplane nose-up direction with the primary surface loaded to correspond to the least of the following airplane nose-up control forces: (1) The maximum control forces expected in service as specified in §§ 25.301 and 25.397. (2) The control force required to produce 1.5 g. (3) The control force corresponding to buffeting or other phenomena of such intensity that it is a strong deterrent to further application of primary longitudinal control force." 14:14:1.0.1.3.13.3.82.1,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.301 Loads.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]","(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads. (b) Unless otherwise provided, the specified air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. Methods used to determine load intensities and distribution must be validated by flight load measurement unless the methods used for determining those loading conditions are shown to be reliable. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account." 14:14:1.0.1.3.13.3.82.2,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.302 Interaction of systems and structures.,FAA,,,"[Doc. No. FAA-2022-1544, 89 FR 68732, Aug. 27, 2024]","For airplanes equipped with systems that affect structural performance, either directly or as a result of a failure or malfunction, the influence of these systems and their failure conditions must be taken into account when showing compliance with the requirements of subparts C and D of this part. These criteria are only applicable to structure whose failure could prevent continued safe flight and landing. (a) General. The applicant must use the following criteria in determining the influence of a system and its failure conditions on the airplane structure. (b) System fully operative. With the system fully operative, the following criteria apply: (1) The applicant must derive limit loads for the limit conditions specified in subpart C of this part, taking into account the behavior of the system up to the limit loads. System nonlinearities must be taken into account. (2) The applicant must show that the airplane meets the strength requirements of subparts C and D of this part, using the appropriate factor of safety to derive ultimate loads from the limit loads defined in paragraph (b)(1) of this section. The effect of nonlinearities must be investigated sufficiently beyond limit conditions to ensure the behavior of the system presents no detrimental effects compared to the behavior below limit conditions. However, conditions beyond limit conditions need not be considered when it can be shown that the airplane has design features that will not allow it to exceed those limit conditions. (3) [Reserved] (c) System in the failure condition. For any system failure condition not shown to be extremely improbable or that results from a single failure, the following criteria apply: (1) At the time of occurrence. The applicant must establish a realistic scenario, starting from 1g level flight conditions, and including pilot corrective actions, to determine the loads occurring at the time of failure and immediately after failure. (i) For static strength substantiation, the airplane must be able to withstand the ultimate loads determined by multiplying the loads in paragraph (c)(1) of this section by a factor of safety that is related to the probability of occurrence of the failure. The factor of safety (F.S.) is defined in Figure 1. (ii) For residual strength substantiation, the airplane must be able to withstand two thirds of the ultimate loads defined in paragraph (c)(1)(i) of this section. For pressurized cabins, these loads must be combined with the normal operating differential pressure. (iii) [Reserved] (iv) Failures of the system that result in forced structural vibrations (oscillatory failures) must not produce loads that could result in detrimental deformation of primary structure. (2) For the continuation of the flight. For the airplane, in the system failed state and considering any appropriate reconfiguration and flight limitations, the following apply: (i) The loads derived from the following conditions at speeds up to V C /M C , or the speed limitation prescribed for the remainder of the flight must be determined: (A) the limit symmetrical maneuvering conditions specified in §§ 25.331 and 25.345, (B) the limit gust and turbulence conditions specified in §§ 25.341 and 25.345, (C) the limit rolling conditions specified in § 25.349 and the limit unsymmetrical conditions specified in §§ 25.367 and 25.427(b) and (c), (D) the limit yaw maneuvering conditions specified in § 25.351, (E) the limit ground loading conditions specified in §§ 25.473 and 25.491, and (F) any other subpart C of this part load condition for which a system is specifically installed or tailored to reduce the loads of that condition. (ii) For static strength substantiation, each part of the structure must be able to withstand the loads in paragraph (c)(2)(i) of this section multiplied by a factor of safety that depends on the probability of being in this failure condition. The factor of safety is defined in Figure 2. Qj = (Tj)(Pj) where: Tj = Average time spent in failure condition j (in hours) Pj = Probability of occurrence of failure mode j (per hour) If Pj is greater than 10 −3 per flight hour, then a 1.5 factor of safety must be applied in lieu of the factor of safety defined in Figure 2. Qj = (Tj)(Pj) where: Tj = Average time spent in failure condition j (in hours) Pj = Probability of occurrence of failure mode j (per hour) If Pj is greater than 10 −3 per flight hour, then a 1.5 factor of safety must be applied in lieu of the factor of safety defined in Figure 2. (iii) For residual strength substantiation, the airplane must be able to withstand two thirds of the ultimate loads defined in paragraph (c)(2)(ii) of this section. For pressurized cabins, these loads must be combined with the normal operating differential pressure. (iv) If the loads induced by the failure condition have a significant effect on fatigue or damage tolerance then their effects must be taken into account. (v)-(vi) [Reserved] (3) [Reserved] (d) Failure indications. For system failure detection and indication, the following apply: (1) The system must be checked for failure conditions evaluated under paragraph (c) of this section that degrade the structural capability below the level required by subparts C (excluding § 25.302) and D of this part or that reduce the reliability of the remaining system. As far as practicable, these failures must be indicated to the flightcrew before flight. (2) The existence of any failure condition evaluated under paragraph (c) of this section that results in a factor of safety between the airplane strength and the loads of subpart C of this part below 1.25 must be indicated to the flightcrew. (e) Dispatch with known failure conditions. If the airplane is to be dispatched in a known system failure condition that affects structural performance or affects the reliability of the remaining system to maintain structural performance, then the Master Minimum Equipment List must ensure the provisions of § 25.302 are met for the dispatched condition and for any subsequent failures. Flight limitations and operational limitations may be taken into account in establishing Qj as the combined probability of being in the dispatched failure condition and the subsequent failure condition for the safety margins in Figure 2. No reduction in these safety margins is allowed if the subsequent system failure rate is greater than 10 −3 per flight hour." 14:14:1.0.1.3.13.3.82.3,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.303 Factor of safety.,FAA,,,"[Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]","Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit load which are considered external loads on the structure. When a loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified." 14:14:1.0.1.3.13.3.82.4,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.305 Strength and deformation.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-77, 57 FR 28949, June 29, 1992; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996]","(a) The structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation. (b) The structure must be able to support ultimate loads without failure for at least 3 seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the 3-second limit does not apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading. When analytical methods are used to show compliance with the ultimate load strength requirements, it must be shown that— (1) The effects of deformation are not significant; (2) The deformations involved are fully accounted for in the analysis; or (3) The methods and assumptions used are sufficient to cover the effects of these deformations. (c) Where structural flexibility is such that any rate of load application likely to occur in the operating conditions might produce transient stresses appreciably higher than those corresponding to static loads, the effects of this rate of application must be considered. (d) [Reserved] (e) The airplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to V D /M D , including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope. This must be shown by analysis, flight tests, or other tests found necessary by the Administrator. (f) Unless shown to be extremely improbable, the airplane must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These must be considered limit loads and must be investigated at airspeeds up to V C /M C ." 14:14:1.0.1.3.13.3.82.5,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.307 Proof of structure.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 25-139, 79 FR 59429, Oct. 2, 2014]","(a) Compliance with the strength and deformation requirements of this subpart must be shown for each critical loading condition. Structural analysis may be used only if the structure conforms to that for which experience has shown this method to be reliable. In other cases, substantiating tests must be made to load levels that are sufficient to verify structural behavior up to loads specified in § 25.305. (b)-(c) [Reserved] (d) When static or dynamic tests are used to show compliance with the requirements of § 25.305(b) for flight structures, appropriate material correction factors must be applied to the test results, unless the structure, or part thereof, being tested has features such that a number of elements contribute to the total strength of the structure and the failure of one element results in the redistribution of the load through alternate load paths." 14:14:1.0.1.3.13.3.83.6,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.321 General.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996]","(a) Flight load factors represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the airplane) to the weight of the airplane. A positive load factor is one in which the aerodynamic force acts upward with respect to the airplane. (b) Considering compressibility effects at each speed, compliance with the flight load requirements of this subpart must be shown— (1) At each critical altitude within the range of altitudes selected by the applicant; (2) At each weight from the design minimum weight to the design maximum weight appropriate to each particular flight load condition; and (3) For each required altitude and weight, for any practicable distribution of disposable load within the operating limitations recorded in the Airplane Flight Manual. (c) Enough points on and within the boundaries of the design envelope must be investigated to ensure that the maximum load for each part of the airplane structure is obtained. (d) The significant forces acting on the airplane must be placed in equilibrium in a rational or conservative manner. The linear inertia forces must be considered in equilibrium with the thrust and all aerodynamic loads, while the angular (pitching) inertia forces must be considered in equilibrium with thrust and all aerodynamic moments, including moments due to loads on components such as tail surfaces and nacelles. Critical thrust values in the range from zero to maximum continuous thrust must be considered." 14:14:1.0.1.3.13.3.84.10,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.337 Limit maneuvering load factors.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]","(a) Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the limit maneuvering load factors prescribed in this section. Pitching velocities appropriate to the corresponding pull-up and steady turn maneuvers must be taken into account. (b) The positive limit maneuvering load factor n for any speed up to Vn may not be less than 2.1 + 24,000/ ( W + 10,000) except that n may not be less than 2.5 and need not be greater than 3.8—where W is the design maximum takeoff weight. (c) The negative limit maneuvering load factor— (1) May not be less than −1.0 at speeds up to V C ; and (2) Must vary linearly with speed from the value at V C to zero at V D . (d) Maneuvering load factors lower than those specified in this section may be used if the airplane has design features that make it impossible to exceed these values in flight." 14:14:1.0.1.3.13.3.84.11,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.341 Gust and turbulence loads.,FAA,,,"[Doc. No. 27902, 61 FR 5221, Feb. 9, 1996; 61 FR 9533, Mar. 8, 1996; Doc. No. FAA-2013-0142; 79 FR 73467, Dec. 11, 2014; Amdt. 25-141, 80 FR 4762, Jan. 29, 2015; 80 FR 6435, Feb. 5, 2015]","(a) Discrete Gust Design Criteria. The airplane is assumed to be subjected to symmetrical vertical and lateral gusts in level flight. Limit gust loads must be determined in accordance with the provisions: (1) Loads on each part of the structure must be determined by dynamic analysis. The analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions. (2) The shape of the gust must be: for 0 ≤s ≤2H where— s = distance penetrated into the gust (feet); U ds = the design gust velocity in equivalent airspeed specified in paragraph (a)(4) of this section; and H = the gust gradient which is the distance (feet) parallel to the airplane's flight path for the gust to reach its peak velocity. where— s = distance penetrated into the gust (feet); U ds = the design gust velocity in equivalent airspeed specified in paragraph (a)(4) of this section; and H = the gust gradient which is the distance (feet) parallel to the airplane's flight path for the gust to reach its peak velocity. (3) A sufficient number of gust gradient distances in the range 30 feet to 350 feet must be investigated to find the critical response for each load quantity. (4) The design gust velocity must be: where— U ref = the reference gust velocity in equivalent airspeed defined in paragraph (a)(5) of this section. F g = the flight profile alleviation factor defined in paragraph (a)(6) of this section. where— U ref = the reference gust velocity in equivalent airspeed defined in paragraph (a)(5) of this section. F g = the flight profile alleviation factor defined in paragraph (a)(6) of this section. (5) The following reference gust velocities apply: (i) At airplane speeds between V B and V C : Positive and negative gusts with reference gust velocities of 56.0 ft/sec EAS must be considered at sea level. The reference gust velocity may be reduced linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS at 15,000 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at 15,000 feet to 20.86 ft/sec EAS at 60,000 feet. (ii) At the airplane design speed V D : The reference gust velocity must be 0.5 times the value obtained under § 25.341(a)(5)(i). (6) The flight profile alleviation factor, F g , must be increased linearly from the sea level value to a value of 1.0 at the maximum operating altitude defined in § 25.1527. At sea level, the flight profile alleviation factor is determined by the following equation: Z mo = Maximum operating altitude defined in § 25.1527 (feet). Z mo = Maximum operating altitude defined in § 25.1527 (feet). (7) When a stability augmentation system is included in the analysis, the effect of any significant system nonlinearities should be accounted for when deriving limit loads from limit gust conditions. (b) Continuous turbulence design criteria. The dynamic response of the airplane to vertical and lateral continuous turbulence must be taken into account. The dynamic analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions. The limit loads must be determined for all critical altitudes, weights, and weight distributions as specified in § 25.321(b), and all critical speeds within the ranges indicated in § 25.341(b)(3). (1) Except as provided in paragraphs (b)(4) and (5) of this section, the following equation must be used: P L = P L−1 g ± U σ A Where— P L = limit load; P L−1g = steady 1g load for the condition; A = ratio of root-mean-square incremental load for the condition to root-mean-square turbulence velocity; and U σ = limit turbulence intensity in true airspeed, specified in paragraph (b)(3) of this section. Where— P L = limit load; P L−1g = steady 1g load for the condition; A = ratio of root-mean-square incremental load for the condition to root-mean-square turbulence velocity; and U σ = limit turbulence intensity in true airspeed, specified in paragraph (b)(3) of this section. (2) Values of A must be determined according to the following formula: Where— H(Ω) = the frequency response function, determined by dynamic analysis, that relates the loads in the aircraft structure to the atmospheric turbulence; and Φ(Ω) = normalized power spectral density of atmospheric turbulence given by— Where— H(Ω) = the frequency response function, determined by dynamic analysis, that relates the loads in the aircraft structure to the atmospheric turbulence; and Φ(Ω) = normalized power spectral density of atmospheric turbulence given by— Where— Ω = reduced frequency, radians per foot; and L = scale of turbulence = 2,500 ft. Where— Ω = reduced frequency, radians per foot; and L = scale of turbulence = 2,500 ft. (3) The limit turbulence intensities, U σ , in feet per second true airspeed required for compliance with this paragraph are— (i) At airplane speeds between V B and V C : U σ = U σ ref F g Where— U σ ref is the reference turbulence intensity that varies linearly with altitude from 90 fps (TAS) at sea level to 79 fps (TAS) at 24,000 feet and is then constant at 79 fps (TAS) up to the altitude of 60,000 feet. F g is the flight profile alleviation factor defined in paragraph (a)(6) of this section; Where— U σ ref is the reference turbulence intensity that varies linearly with altitude from 90 fps (TAS) at sea level to 79 fps (TAS) at 24,000 feet and is then constant at 79 fps (TAS) up to the altitude of 60,000 feet. F g is the flight profile alleviation factor defined in paragraph (a)(6) of this section; (ii) At speed V D : U σ is equal to 1/2 the values obtained under paragraph (b)(3)(i) of this section. (iii) At speeds between V C and V D : U σ is equal to a value obtained by linear interpolation. (iv) At all speeds, both positive and negative incremental loads due to continuous turbulence must be considered. (4) When an automatic system affecting the dynamic response of the airplane is included in the analysis, the effects of system non-linearities on loads at the limit load level must be taken into account in a realistic or conservative manner. (5) If necessary for the assessment of loads on airplanes with significant non-linearities, it must be assumed that the turbulence field has a root-mean-square velocity equal to 40 percent of the U σ values specified in paragraph (b)(3) of this section. The value of limit load is that load with the same probability of exceedance in the turbulence field as A U σ of the same load quantity in a linear approximated model. (c) Supplementary gust conditions for wing-mounted engines. For airplanes equipped with wing-mounted engines, the engine mounts, pylons, and wing supporting structure must be designed for the maximum response at the nacelle center of gravity derived from the following dynamic gust conditions applied to the airplane: (1) A discrete gust determined in accordance with § 25.341(a) at each angle normal to the flight path, and separately, (2) A pair of discrete gusts, one vertical and one lateral. The length of each of these gusts must be independently tuned to the maximum response in accordance with § 25.341(a). The penetration of the airplane in the combined gust field and the phasing of the vertical and lateral component gusts must be established to develop the maximum response to the gust pair. In the absence of a more rational analysis, the following formula must be used for each of the maximum engine loads in all six degrees of freedom: Where— P L = limit load; P L-1g = steady 1g load for the condition; L V = peak incremental response load due to a vertical gust according to § 25.341(a); and L L = peak incremental response load due to a lateral gust according to § 25.341(a). Where— P L = limit load; P L-1g = steady 1g load for the condition; L V = peak incremental response load due to a vertical gust according to § 25.341(a); and L L = peak incremental response load due to a lateral gust according to § 25.341(a)." 14:14:1.0.1.3.13.3.84.12,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.343 Design fuel and oil loads.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 33 FR 12226, Aug. 30, 1968; Amdt. 25-72, 55 FR 37607, Sept. 12, 1990; Amdt. 25-86, 61 FR 5221, Feb. 9, 1996; Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) The disposable load combinations must include each fuel and oil load in the range from zero fuel and oil to the selected maximum fuel and oil load. A structural reserve fuel condition, not exceeding 45 minutes of fuel under the operating conditions in § 25.1001(e) and (f), as applicable, may be selected. (b) If a structural reserve fuel condition is selected, it must be used as the minimum fuel weight condition for showing compliance with the flight load requirements as prescribed in this subpart. In addition— (1) The structure must be designed for a condition of zero fuel and oil in the wing at limit loads corresponding to— (i) A maneuvering load factor of + 2.25; and (ii) The gust and turbulence conditions of § 25.341(a) and (b), but assuming 85% of the gust velocities prescribed in § 25.341(a)(4) and 85% of the turbulence intensities prescribed in § 25.341(b)(3). (2) Fatigue evaluation of the structure must account for any increase in operating stresses resulting from the design condition of paragraph (b)(1) of this section; and (3) The flutter, deformation, and vibration requirements must also be met with zero fuel." 14:14:1.0.1.3.13.3.84.13,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.345 High lift devices.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 37607, Sept. 17, 1990; Amdt. 25-86, 61 FR 5221, Feb. 9, 1996; Amdt. 25-91, 62 FR 40704, July 29, 1997; Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) If wing flaps are to be used during takeoff, approach, or landing, at the design flap speeds established for these stages of flight under § 25.335(e) and with the wing flaps in the corresponding positions, the airplane is assumed to be subjected to symmetrical maneuvers and gusts. The resulting limit loads must correspond to the conditions determined as follows: (1) Maneuvering to a positive limit load factor of 2.0; and (2) Positive and negative gusts of 25 ft/sec EAS acting normal to the flight path in level flight. Gust loads resulting on each part of the structure must be determined by rational analysis. The analysis must take into account the unsteady aerodynamic characteristics and rigid body motions of the aircraft. The shape of the gust must be as described in § 25.341(a)(2) except that— U ds = 25 ft/sec EAS; H = 12.5 c; and c = mean geometric chord of the wing (feet). U ds = 25 ft/sec EAS; H = 12.5 c; and c = mean geometric chord of the wing (feet). (b) The airplane must be designed for the conditions prescribed in paragraph (a) of this section, except that the airplane load factor need not exceed 1.0, taking into account, as separate conditions, the effects of— (1) Propeller slipstream corresponding to maximum continuous power at the design flap speeds V F, and with takeoff power at not less than 1.4 times the stalling speed for the particular flap position and associated maximum weight; and (2) A head-on gust of 25 feet per second velocity (EAS). (c) If flaps or other high lift devices are to be used in en route conditions, and with flaps in the appropriate position at speeds up to the flap design speed chosen for these conditions, the airplane is assumed to be subjected to symmetrical maneuvers and gusts within the range determined by— (1) Maneuvering to a positive limit load factor as prescribed in § 25.337(b); and (2) The vertical gust and turbulence conditions prescribed in § 25.341(a) and (b). (d) The airplane must be designed for a maneuvering load factor of 1.5 g at the maximum take-off weight with the wing-flaps and similar high lift devices in the landing configurations." 14:14:1.0.1.3.13.3.84.14,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.349 Rolling conditions.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5222, Feb. 9, 1996; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998]","The airplane must be designed for loads resulting from the rolling conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces. (a) Maneuvering. The following conditions, speeds, and aileron deflections (except as the deflections may be limited by pilot effort) must be considered in combination with an airplane load factor of zero and of two-thirds of the positive maneuvering factor used in design. In determining the required aileron deflections, the torsional flexibility of the wing must be considered in accordance with § 25.301(b): (1) Conditions corresponding to steady rolling velocities must be investigated. In addition, conditions corresponding to maximum angular acceleration must be investigated for airplanes with engines or other weight concentrations outboard of the fuselage. For the angular acceleration conditions, zero rolling velocity may be assumed in the absence of a rational time history investigation of the maneuver. (2) At V A, a sudden deflection of the aileron to the stop is assumed. (3) At V C, the aileron deflection must be that required to produce a rate of roll not less than that obtained in paragraph (a)(2) of this section. (4) At V D, the aileron deflection must be that required to produce a rate of roll not less than one-third of that in paragraph (a)(2) of this section. (b) Unsymmetrical gusts. The airplane is assumed to be subjected to unsymmetrical vertical gusts in level flight. The resulting limit loads must be determined from either the wing maximum airload derived directly from § 25.341(a), or the wing maximum airload derived indirectly from the vertical load factor calculated from § 25.341(a). It must be assumed that 100 percent of the wing air load acts on one side of the airplane and 80 percent of the wing air load acts on the other side." 14:14:1.0.1.3.13.3.84.15,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.351 Yaw maneuver conditions.,FAA,,,"[Amdt. 25-91, 62 FR 40704, July 29, 1997]","The airplane must be designed for loads resulting from the yaw maneuver conditions specified in paragraphs (a) through (d) of this section at speeds from V MC to V D . Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero. (a) With the airplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by: (1) The control system on control surface stops; or (2) A limit pilot force of 300 pounds from V MC to V A and 200 pounds from V C /M C to V D /M D , with a linear variation between V A and V C /M C . (b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the airplane yaws to the overswing sideslip angle. (c) With the airplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section. (d) With the airplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral." 14:14:1.0.1.3.13.3.84.16,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.353 Rudder control reversal conditions.,FAA,,,"[Amdt. No. 25-147, 87 FR 71210, Nov. 22, 2022]","Airplanes with a powered rudder control surface or surfaces must be designed for loads, considered to be ultimate, resulting from the yaw maneuver conditions specified in paragraphs (a) through (e) of this section at speeds from V MC to V C /M C . Any permanent deformation resulting from these ultimate load conditions must not prevent continued safe flight and landing. The applicant must evaluate these conditions with the landing gear retracted and speed brakes (and spoilers when used as speed brakes) retracted. The applicant must evaluate the effects of flaps, flaperons, or any other aerodynamic devices when used as flaps, and slats-extended configurations, if they are used in en route conditions. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the loads on the airplane, the yawing velocity may be assumed to be zero. The applicant must assume a pilot force of 200 pounds when evaluating each of the following conditions: (a) With the airplane in unaccelerated flight at zero yaw, the flightdeck rudder control is suddenly and fully displaced to achieve the resulting rudder deflection, as limited by the control system or the control surface stops. (b) With the airplane yawed to the overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops. (c) With the airplane yawed to the opposite overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops. (d) With the airplane yawed to the subsequent overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops. (e) With the airplane yawed to the opposite overswing sideslip angle, the flightdeck rudder control is suddenly returned to neutral." 14:14:1.0.1.3.13.3.84.7,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.331 Symmetric maneuvering conditions.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50594, Oct. 30, 1978; 43 FR 52495, Nov. 13, 1978; 43 FR 54082, Nov. 20, 1978; Amdt. 25-72, 55 FR 29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996; Amdt. 25-91, 62 FR 40704, July 29, 1997; Amdt. 25-141, 79 FR 73466, Dec. 11, 2014]","(a) Procedure. For the analysis of the maneuvering flight conditions specified in paragraphs (b) and (c) of this section, the following provisions apply: (1) Where sudden displacement of a control is specified, the assumed rate of control surface displacement may not be less than the rate that could be applied by the pilot through the control system. (2) In determining elevator angles and chordwise load distribution in the maneuvering conditions of paragraphs (b) and (c) of this section, the effect of corresponding pitching velocities must be taken into account. The in-trim and out-of-trim flight conditions specified in § 25.255 must be considered. (b) Maneuvering balanced conditions. Assuming the airplane to be in equilibrium with zero pitching acceleration, the maneuvering conditions A through I on the maneuvering envelope in § 25.333(b) must be investigated. (c) Maneuvering pitching conditions. The following conditions must be investigated: (1) Maximum pitch control displacement at V A . The airplane is assumed to be flying in steady level flight (point A 1 , § 25.333(b)) and the cockpit pitch control is suddenly moved to obtain extreme nose up pitching acceleration. In defining the tail load, the response of the airplane must be taken into account. Airplane loads that occur subsequent to the time when normal acceleration at the c.g. exceeds the positive limit maneuvering load factor (at point A 2 in § 25.333(b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered. (2) Checked maneuver between V A and V D . Nose-up checked pitching maneuvers must be analyzed in which the positive limit load factor prescribed in § 25.337 is achieved. As a separate condition, nose-down checked pitching maneuvers must be analyzed in which a limit load factor of 0g is achieved. In defining the airplane loads, the flight deck pitch control motions described in paragraphs (c)(2)(i) through (iv) of this section must be used: (i) The airplane is assumed to be flying in steady level flight at any speed between V A and V D and the flight deck pitch control is moved in accordance with the following formula: δ(t) = δ 1 sin(ωt) for 0 ≤ t ≤ t max Where— δ 1 = the maximum available displacement of the flight deck pitch control in the initial direction, as limited by the control system stops, control surface stops, or by pilot effort in accordance with § 25.397(b); δ(t) = the displacement of the flight deck pitch control as a function of time. In the initial direction, δ(t) is limited to δ 1 . In the reverse direction, δ(t) may be truncated at the maximum available displacement of the flight deck pitch control as limited by the control system stops, control surface stops, or by pilot effort in accordance with 25.397(b); t max = 3π/2ω; ω = the circular frequency (radians/second) of the control deflection taken equal to the undamped natural frequency of the short period rigid mode of the airplane, with active control system effects included where appropriate; but not less than: Where V = the speed of the airplane at entry to the maneuver. V A = the design maneuvering speed prescribed in § 25.335(c). Where— δ 1 = the maximum available displacement of the flight deck pitch control in the initial direction, as limited by the control system stops, control surface stops, or by pilot effort in accordance with § 25.397(b); δ(t) = the displacement of the flight deck pitch control as a function of time. In the initial direction, δ(t) is limited to δ 1 . In the reverse direction, δ(t) may be truncated at the maximum available displacement of the flight deck pitch control as limited by the control system stops, control surface stops, or by pilot effort in accordance with 25.397(b); t max = 3π/2ω; ω = the circular frequency (radians/second) of the control deflection taken equal to the undamped natural frequency of the short period rigid mode of the airplane, with active control system effects included where appropriate; but not less than: Where V = the speed of the airplane at entry to the maneuver. V A = the design maneuvering speed prescribed in § 25.335(c). (ii) For nose-up pitching maneuvers, the complete flight deck pitch control displacement history may be scaled down in amplitude to the extent necessary to ensure that the positive limit load factor prescribed in § 25.337 is not exceeded. For nose-down pitching maneuvers, the complete flight deck control displacement history may be scaled down in amplitude to the extent necessary to ensure that the normal acceleration at the center of gravity does not go below 0g. (iii) In addition, for cases where the airplane response to the specified flight deck pitch control motion does not achieve the prescribed limit load factors, then the following flight deck pitch control motion must be used: δ(t) = δ 1 sin(ωt) for 0 ≤ t ≤ t 1 δ(t) = δ 1 for t 1 ≤ t ≤ t 2 δ(t) = δ 1 sin(ω[t + t 1 − t 2 ]) for t 2 ≤ t ≤ t max Where— t 1 = π/2ω t 2 = t 1 + Δt t max = t 2 + π/ω; Δt = the minimum period of time necessary to allow the prescribed limit load factor to be achieved in the initial direction, but it need not exceed five seconds (see figure below). Where— t 1 = π/2ω t 2 = t 1 + Δt t max = t 2 + π/ω; Δt = the minimum period of time necessary to allow the prescribed limit load factor to be achieved in the initial direction, but it need not exceed five seconds (see figure below). (iv) In cases where the flight deck pitch control motion may be affected by inputs from systems (for example, by a stick pusher that can operate at high load factor as well as at 1g), then the effects of those systems shall be taken into account. (v) Airplane loads that occur beyond the following times need not be considered: (A) For the nose-up pitching maneuver, the time at which the normal acceleration at the center of gravity goes below 0g; (B) For the nose-down pitching maneuver, the time at which the normal acceleration at the center of gravity goes above the positive limit load factor prescribed in § 25.337; (C) t max. ." 14:14:1.0.1.3.13.3.84.8,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.333 Flight maneuvering envelope.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-86, 61 FR 5220, Feb. 9, 1996]","(a) General. The strength requirements must be met at each combination of airspeed and load factor on and within the boundaries of the representative maneuvering envelope ( V-n diagram) of paragraph (b) of this section. This envelope must also be used in determining the airplane structural operating limitations as specified in § 25.1501. (b) Maneuvering envelope." 14:14:1.0.1.3.13.3.84.9,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.335 Design airspeeds.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-86, 61 FR 5220, Feb. 9, 1996; Amdt. 25-91, 62 FR 40704, July 29, 1997]","The selected design airspeeds are equivalent airspeeds (EAS). Estimated values of V S 0 and V S 1 must be conservative. (a) Design cruising speed, V C . For V C, the following apply: (1) The minimum value of V C must be sufficiently greater than V B to provide for inadvertent speed increases likely to occur as a result of severe atmospheric turbulence. (2) Except as provided in § 25.335(d)(2), V C may not be less than V B + 1.32 U REF (with U REF as specified in § 25.341(a)(5)(i)). However V C need not exceed the maximum speed in level flight at maximum continuous power for the corresponding altitude. (3) At altitudes where V D is limited by Mach number, V C may be limited to a selected Mach number. (b) Design dive speed, V D . V D must be selected so that V C / M C is not greater than 0.8 V D / M D, or so that the minimum speed margin between V C / M C and V D / M D is the greater of the following values: (1) From an initial condition of stabilized flight at V C / M C, the airplane is upset, flown for 20 seconds along a flight path 7.5° below the initial path, and then pulled up at a load factor of 1.5 g (0.5 g acceleration increment). The speed increase occurring in this maneuver may be calculated if reliable or conservative aerodynamic data is used. Power as specified in § 25.175(b)(1)(iv) is assumed until the pullup is initiated, at which time power reduction and the use of pilot controlled drag devices may be assumed; (2) The minimum speed margin must be enough to provide for atmospheric variations (such as horizontal gusts, and penetration of jet streams and cold fronts) and for instrument errors and airframe production variations. These factors may be considered on a probability basis. The margin at altitude where M C is limited by compressibility effects must not less than 0.07M unless a lower margin is determined using a rational analysis that includes the effects of any automatic systems. In any case, the margin may not be reduced to less than 0.05M. (c) Design maneuvering speed V A . For V A , the following apply: (1) V A may not be less than V S 1 √n where— (i) n is the limit positive maneuvering load factor at V C ; and (ii) V S 1 is the stalling speed with flaps retracted. (2) V A and V S must be evaluated at the design weight and altitude under consideration. (3) V A need not be more than V C or the speed at which the positive C N max curve intersects the positive maneuver load factor line, whichever is less. (d) Design speed for maximum gust intensity, V B . (1) V B may not be less than where— V S1 = the 1-g stalling speed based on C NAmax with the flaps retracted at the particular weight under consideration; V c = design cruise speed (knots equivalent airspeed); U ref = the reference gust velocity (feet per second equivalent airspeed) from § 25.341(a)(5)(i); w = average wing loading (pounds per square foot) at the particular weight under consideration. where— V S1 = the 1-g stalling speed based on C NAmax with the flaps retracted at the particular weight under consideration; V c = design cruise speed (knots equivalent airspeed); U ref = the reference gust velocity (feet per second equivalent airspeed) from § 25.341(a)(5)(i); w = average wing loading (pounds per square foot) at the particular weight under consideration. ρ = density of air (slugs/ft 3 ); c = mean geometric chord of the wing (feet); g = acceleration due to gravity (ft/sec 2 ); a = slope of the airplane normal force coefficient curve, C NA per radian; ρ = density of air (slugs/ft 3 ); c = mean geometric chord of the wing (feet); g = acceleration due to gravity (ft/sec 2 ); a = slope of the airplane normal force coefficient curve, C NA per radian; (2) At altitudes where V C is limited by Mach number— (i) V B may be chosen to provide an optimum margin between low and high speed buffet boundaries; and, (ii) V B need not be greater than V C . (e) Design flap speeds, V F . For V F , the following apply: (1) The design flap speed for each flap position (established in accordance with § 25.697(a)) must be sufficiently greater than the operating speed recommended for the corresponding stage of flight (including balked landings) to allow for probable variations in control of airspeed and for transition from one flap position to another. (2) If an automatic flap positioning or load limiting device is used, the speeds and corresponding flap positions programmed or allowed by the device may be used. (3) V F may not be less than— (i) 1.6 V S 1 with the flaps in takeoff position at maximum takeoff weight; (ii) 1.8 V S 1 with the flaps in approach position at maximum landing weight, and (iii) 1.8 V S 0 with the flaps in landing position at maximum landing weight. (f) Design drag device speeds, V DD . The selected design speed for each drag device must be sufficiently greater than the speed recommended for the operation of the device to allow for probable variations in speed control. For drag devices intended for use in high speed descents, V DD may not be less than V D . When an automatic drag device positioning or load limiting means is used, the speeds and corresponding drag device positions programmed or allowed by the automatic means must be used for design." 14:14:1.0.1.3.13.3.85.17,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.361 Engine and auxiliary power unit torque.,FAA,,,"[Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) For engine installations— (1) Each engine mount, pylon, and adjacent supporting airframe structures must be designed for the effects of— (i) A limit engine torque corresponding to takeoff power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with 75% of the limit loads from flight condition A of § 25.333(b); (ii) A limit engine torque corresponding to the maximum continuous power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with the limit loads from flight condition A of § 25.333(b); and (iii) For turbopropeller installations only, in addition to the conditions specified in paragraphs (a)(1)(i) and (ii) of this section, a limit engine torque corresponding to takeoff power and propeller speed, multiplied by a factor accounting for propeller control system malfunction, including quick feathering, acting simultaneously with 1g level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used. (2) The limit engine torque to be considered under paragraph (a)(1) of this section must be obtained by— (i) For turbopropeller installations, multiplying mean engine torque for the specified power/thrust and speed by a factor of 1.25; (ii) For other turbine engines, the limit engine torque must be equal to the maximum accelerating torque for the case considered. (3) The engine mounts, pylons, and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit engine torque loads imposed by each of the following conditions to be considered separately: (i) Sudden maximum engine deceleration due to malfunction or abnormal condition; and (ii) The maximum acceleration of engine. (b) For auxiliary power unit installations, the power unit mounts and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit torque loads imposed by each of the following conditions to be considered separately: (1) Sudden maximum auxiliary power unit deceleration due to malfunction, abnormal condition, or structural failure; and (2) The maximum acceleration of the auxiliary power unit." 14:14:1.0.1.3.13.3.85.18,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.362 Engine failure loads.,FAA,,,"[Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) For engine mounts, pylons, and adjacent supporting airframe structure, an ultimate loading condition must be considered that combines 1g flight loads with the most critical transient dynamic loads and vibrations, as determined by dynamic analysis, resulting from failure of a blade, shaft, bearing or bearing support, or bird strike event. Any permanent deformation from these ultimate load conditions must not prevent continued safe flight and landing. (b) The ultimate loads developed from the conditions specified in paragraph (a) of this section are to be— (1) Multiplied by a factor of 1.0 when applied to engine mounts and pylons; and (2) Multiplied by a factor of 1.25 when applied to adjacent supporting airframe structure." 14:14:1.0.1.3.13.3.85.19,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.363 Side load on engine and auxiliary power unit mounts.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-91, 62 FR 40704, July 29, 1997]","(a) Each engine and auxiliary power unit mount and its supporting structure must be designed for a limit load factor in lateral direction, for the side load on the engine and auxiliary power unit mount, at least equal to the maximum load factor obtained in the yawing conditions but not less than— (1) 1.33; or (2) One-third of the limit load factor for flight condition A as prescribed in § 25.333(b). (b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions." 14:14:1.0.1.3.13.3.85.20,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.365 Pressurized compartment loads.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-71, 55 FR 13477, Apr. 10, 1990; Amdt. 25-72, 55 FR 29776, July 20, 1990; Amdt. 25-87, 61 FR 28695, June 5, 1996; Amdt. No. 25-149, 88 FR 38382, June 13, 2023]","For airplanes with one or more pressurized compartments the following apply: (a) The airplane structure must be strong enough to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting. (b) The external pressure distribution in flight, and stress concentrations and fatigue effects must be accounted for. (c) If landings may be made with the compartment pressurized, landing loads must be combined with pressure differential loads from zero up to the maximum allowed during landing. (d) The airplane structure must be designed to be able to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33 for airplanes to be approved for operation to 45,000 feet or by a factor of 1.67 for airplanes to be approved for operation above 45,000 feet, omitting other loads. (e) Any structure, component or part, inside or outside a pressurized compartment, the failure of which could interfere with continued safe flight and landing, must be designed to withstand the effects of a sudden release of pressure through an opening in any compartment at any operating altitude resulting from each of the following conditions: (1) The penetration of the compartment by a portion of an engine following an engine disintegration; (2) Any opening in any pressurized compartment up to the size H o in square feet; however, small compartments may be combined with an adjacent pressurized compartment and both considered as a single compartment for openings that cannot reasonably be expected to be confined to the small compartment. The size H o must be computed by the following formula: H o = PA s where, H o = Maximum opening in square feet, need not exceed 20 square feet. P = (A s /6240) + .024 A s = Maximum cross-sectional area of the pressurized shell normal to the longitudinal axis, in square feet; and where, H o = Maximum opening in square feet, need not exceed 20 square feet. P = (A s /6240) + .024 A s = Maximum cross-sectional area of the pressurized shell normal to the longitudinal axis, in square feet; and (3) The maximum opening caused by airplane or equipment failures not shown to be extremely improbable. (f) In complying with paragraph (e) of this section, the fail-safe features of the design may be considered in determining the probability of failure or penetration and probable size of openings, provided that possible improper operation of closure devices and inadvertent door openings are also considered. Furthermore, the resulting differential pressure loads must be combined in a rational and conservative manner with 1-g level flight loads and any loads arising from emergency depressurization conditions. These loads may be considered as ultimate conditions; however, any deformations associated with these conditions must not interfere with continued safe flight and landing. The pressure relief provided by intercompartment venting may also be considered. (g)(1) Except as provided in paragraph (g)(2) of this section, bulkheads, floors, and partitions in pressurized compartments for occupants must be designed to withstand the conditions specified in paragraph (e) of this section. In addition, reasonable design precautions must be taken to minimize the probability of parts becoming detached and injuring occupants while in their seats. (2) Partitions adjacent to the opening specified in paragraph (e)(2) of this section need not be designed to withstand that condition provided— (i) Failure of the partition would not interfere with continued safe flight and landing; and (ii) Designing the partition to withstand the condition specified in paragraph (e)(2) of this section would be impractical." 14:14:1.0.1.3.13.3.85.21,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.367 Unsymmetrical loads due to engine failure.,FAA,,,,"(a) The airplane must be designed for the unsymmetrical loads resulting from the failure of the critical engine. Turbopropeller airplanes must be designed for the following conditions in combination with a single malfunction of the propeller drag limiting system, considering the probable pilot corrective action on the flight controls: (1) At speeds between V MC and V D, the loads resulting from power failure because of fuel flow interruption are considered to be limit loads. (2) At speeds between V MC and V C, the loads resulting from the disconnection of the engine compressor from the turbine or from loss of the turbine blades are considered to be ultimate loads. (3) The time history of the thrust decay and drag build-up occurring as a result of the prescribed engine failures must be substantiated by test or other data applicable to the particular engine-propeller combination. (4) The timing and magnitude of the probable pilot corrective action must be conservatively estimated, considering the characteristics of the particular engine-propeller-airplane combination. (b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than two seconds after the engine failure. The magnitude of the corrective action may be based on the control forces specified in § 25.397(b) except that lower forces may be assumed where it is shown by anaylsis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions." 14:14:1.0.1.3.13.3.85.22,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.371 Gyroscopic loads.,FAA,,,"[Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","The structure supporting any engine or auxiliary power unit must be designed for the loads, including gyroscopic loads, arising from the conditions specified in §§ 25.331, 25.341, 25.349, 25.351, 25.473, 25.479, and 25.481, with the engine or auxiliary power unit at the maximum rotating speed appropriate to the condition. For the purposes of compliance with this paragraph, the pitch maneuver in § 25.331(c)(1) must be carried out until the positive limit maneuvering load factor (point A 2 in § 25.333(b)) is reached." 14:14:1.0.1.3.13.3.85.23,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.373 Speed control devices.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR 29776, July 20, 1990; Amdt. 25-86, 61 FR 5222, Feb. 9, 1996; Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","If speed control devices (such as spoilers and drag flaps) are installed for use in en route conditions— (a) The airplane must be designed for the symmetrical maneuvers prescribed in §§ 25.333 and 25.337, the yawing maneuvers in § 25.351, and the vertical and lateral gust and turbulence conditions prescribed in § 25.341(a) and (b) at each setting and the maximum speed associated with that setting; and (b) If the device has automatic operating or load limiting features, the airplane must be designed for the maneuver and gust conditions prescribed in paragraph (a) of this section, at the speeds and corresponding device positions that the mechanism allows." 14:14:1.0.1.3.13.3.86.24,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.391 Control surface loads: General.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-86, 61 FR 5222, Feb. 9, 1996; Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","The control surfaces must be designed for the limit loads resulting from the flight conditions in §§ 25.331, 25.341(a) and (b), 25.349, and 25.351, considering the requirements for— (a) Loads parallel to hinge line, in § 25.393; (b) Pilot effort effects, in § 25.397; (c) Trim tab effects, in § 25.407; (d) Unsymmetrical loads, in § 25.427; and (e) Auxiliary aerodynamic surfaces, in § 25.445." 14:14:1.0.1.3.13.3.86.25,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.393 Loads parallel to hinge line.,FAA,,,,"(a) Control surfaces and supporting hinge brackets must be designed for inertia loads acting parallel to the hinge line. (b) In the absence of more rational data, the inertia loads may be assumed to be equal to KW, where— (1) K = 24 for vertical surfaces; (2) K = 12 for horizontal surfaces; and (3) W = weight of the movable surfaces." 14:14:1.0.1.3.13.3.86.26,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.395 Control system.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5672, Apr. 8, 1970; Amdt. 25-72, 55 FR 29776, July 20, 1990; Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) Longitudinal, lateral, directional, and drag control system and their supporting structures must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in § 25.391. (b) The system limit loads of paragraph (a) of this section need not exceed the loads that can be produced by the pilot (or pilots) and by automatic or power devices operating the controls. (c) The loads must not be less than those resulting from application of the minimum forces prescribed in § 25.397(c)." 14:14:1.0.1.3.13.3.86.27,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.397 Control system loads.,FAA,,,"[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 25-72, 55 FR 29776, July 20, 1990]","(a) General. The maximum and minimum pilot forces, specified in paragraph (c) of this section, are assumed to act at the appropriate control grips or pads (in a manner simulating flight conditions) and to be reacted at the attachment of the control system to the control surface horn. (b) Pilot effort effects. In the control surface flight loading condition, the air loads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (c) of this section. Two-thirds of the maximum values specified for the aileron and elevator may be used if control surface hinge moments are based on reliable data. In applying this criterion, the effects of servo mechanisms, tabs, and automatic pilot systems, must be considered. (c) Limit pilot forces and torques. The limit pilot forces and torques are as follows: 1 The critical parts of the aileron control system must be designed for a single tangential force with a limit value equal to 1.25 times the couple force determined from these criteria. 2 D = wheel diameter (inches). 3 The unsymmetrical forces must be applied at one of the normal handgrip points on the periphery of the control wheel." 14:14:1.0.1.3.13.3.86.28,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.399 Dual control system.,FAA,,,,"(a) Each dual control system must be designed for the pilots operating in opposition, using individual pilot forces not less than— (1) 0.75 times those obtained under § 25.395; or (2) The minimum forces specified in § 25.397(c). (b) The control system must be designed for pilot forces applied in the same direction, using individual pilot forces not less than 0.75 times those obtained under § 25.395." 14:14:1.0.1.3.13.3.86.29,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.405 Secondary control system.,FAA,,,,"Secondary controls, such as wheel brake, spoiler, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls. The following values may be used: Pilot Control Force Limits (Secondary Controls) *Limited to flap, tab, stabilizer, spoiler, and landing gear operation controls." 14:14:1.0.1.3.13.3.86.30,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.407 Trim tab effects.,FAA,,,,"The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort. In these cases, the tabs are considered to be deflected in the direction that would assist the pilot, and the deflections are— (a) For elevator trim tabs, those required to trim the airplane at any point within the positive portion of the pertinent flight envelope in § 25.333(b), except as limited by the stops; and (b) For aileron and rudder trim tabs, those required to trim the airplane in the critical unsymmetrical power and loading conditions, with appropriate allowance for rigging tolerances." 14:14:1.0.1.3.13.3.86.31,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.409 Tabs.,FAA,,,,"(a) Trim tabs. Trim tabs must be designed to withstand loads arising from all likely combinations of tab setting, primary control position, and airplane speed (obtainable without exceeding the flight load conditions prescribed for the airplane as a whole), when the effect of the tab is opposed by pilot effort forces up to those specified in § 25.397(b). (b) Balancing tabs. Balancing tabs must be designed for deflections consistent with the primary control surface loading conditions. (c) Servo tabs. Servo tabs must be designed for deflections consistent with the primary control surface loading conditions obtainable within the pilot maneuvering effort, considering possible opposition from the trim tabs." 14:14:1.0.1.3.13.3.86.32,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.415 Ground gust conditions.,FAA,,,"[Amdt. 25-141, 79 FR 73468, Dec. 11, 2014]","(a) The flight control systems and surfaces must be designed for the limit loads generated when the airplane is subjected to a horizontal 65-knot ground gust from any direction while taxiing and while parked. For airplanes equipped with control system gust locks, the taxiing condition must be evaluated with the controls locked and unlocked, and the parked condition must be evaluated with the controls locked. (b) The control system and surface loads due to ground gust may be assumed to be static loads, and the hinge moments H must be computed from the formula: H = K (1/2) ρ o V 2 c S Where— K = hinge moment factor for ground gusts derived in paragraph (c) of this section; ρ o = density of air at sea level; V = 65 knots relative to the aircraft; S = area of the control surface aft of the hinge line; c = mean aerodynamic chord of the control surface aft of the hinge line. Where— K = hinge moment factor for ground gusts derived in paragraph (c) of this section; ρ o = density of air at sea level; V = 65 knots relative to the aircraft; S = area of the control surface aft of the hinge line; c = mean aerodynamic chord of the control surface aft of the hinge line. (c) The hinge moment factor K for ground gusts must be taken from the following table: * A positive value of K indicates a moment tending to depress the surface, while a negative value of K indicates a moment tending to raise the surface. (d) The computed hinge moment of paragraph (b) of this section must be used to determine the limit loads due to ground gust conditions for the control surface. A 1.25 factor on the computed hinge moments must be used in calculating limit control system loads. (e) Where control system flexibility is such that the rate of load application in the ground gust conditions might produce transient stresses appreciably higher than those corresponding to static loads, in the absence of a rational analysis substantiating a different dynamic factor, an additional factor of 1.6 must be applied to the control system loads of paragraph (d) of this section to obtain limit loads. If a rational analysis is used, the additional factor must not be less than 1.2. (f) For the condition of the control locks engaged, the control surfaces, the control system locks, and the parts of any control systems between the surfaces and the locks must be designed to the resultant limit loads. Where control locks are not provided, then the control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads. If the control system design is such as to allow any part of the control system to impact with the stops due to flexibility, then the resultant impact loads must be taken into account in deriving the limit loads due to ground gust. (g) For the condition of taxiing with the control locks disengaged, or where control locks are not provided, the following apply: (1) The control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads. (2) The parts of the control systems between the stops nearest the surfaces and the flight deck controls must be designed to the resultant limit loads, except that the parts of the control system where loads are eventually reacted by the pilot need not exceed: (i) The loads corresponding to the maximum pilot loads in § 25.397(c) for each pilot alone; or (ii) 0.75 times these maximum loads for each pilot when the pilot forces are applied in the same direction." 14:14:1.0.1.3.13.3.86.33,14,Aeronautics and Space,I,C,25,PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES,C,Subpart C—Structure,,§ 25.427 Unsymmetrical loads.,FAA,,,"[Doc. No. 27902, 61 FR 5222, Feb. 9, 1996]","(a) In designing the airplane for lateral gust, yaw maneuver and roll maneuver conditions, account must be taken of unsymmetrical loads on the empennage arising from effects such as slipstream and aerodynamic interference with the wing, vertical fin and other aerodynamic surfaces. (b) The horizontal tail must be assumed to be subjected to unsymmetrical loading conditions determined as follows: (1) 100 percent of the maximum loading from the symmetrical maneuver conditions of § 25.331 and the vertical gust conditions of § 25.341(a) acting separately on the surface on one side of the plane of symmetry; and (2) 80 percent of these loadings acting on the other side. (c) For empennage arrangements where the horizontal tail surfaces have dihedral angles greater than plus or minus 10 degrees, or are supported by the vertical tail surfaces, the surfaces and the supporting structure must be designed for gust velocities specified in § 25.341(a) acting in any orientation at right angles to the flight path. (d) Unsymmetrical loading on the empennage arising from buffet conditions of § 25.305(e) must be taken into account."