(iv) Load the pilot’s control until the control surface is just off the support.
(v) Determine the available control surface travel, which is the amount of movement of the surface from neutral when the cockpit control is moved through the limits of its travel.
(vi) The control surface under loads described in paragraph f above must have adequate flight characteristics as specified in § 23.141.
(vii) To address the possibility of a critical intermediate control surface loading, gradually remove load from the control surface (while maintaining the load on the adjacent fixed surface) until maximum control surface travel is achieved.
(viii) The above procedure should be repeated in the opposite direction.
(ix) With limit load applied to the adjacent fixed surface and limit or intermediate load applied to the control surface, no signs of jamming, or of any permanent set of any connection, bracket, attachment, and so forth, may be present.
(x) The control system should operate freely without excessive friction.
Excessive friction is any increase under limit loads that results in exceeding the limit control forces and torques specified by the regulations.
(xi) Cable systems should be checked with the loads applied to ensure that excessive slack does not develop in the system. Excessive slack is any change in cables or cable hardware that results in reduced airplane control surface movement.
(xii) Repeat this process for each of the critical loading conditions as defined by paragraphs d and f above.
SECTION (2):
(2) The tests described in this section support the demonstration that structural deformations not interfere with safe operation as required by § 23.305, paragraph
(a). Accomplish the following:
(i) Load the adjacent fixed aerodynamic surface (wing, horizontal tail, or vertical tail) in accordance with one of the conditions of paragraph d and e above.
(ii) Operate the unloaded control system from stop to stop.
(iii) No signs of interference (contact) may be present.
(iv) The control system should operate freely without excessive friction.
(v) Repeat this process for each of the critical adjacent fixed surface loading conditions as defined by paragraphs d and e above.
NOTE 1: An alternate procedure may be used to accommodate the testing described in sections (1) and (2) above during structural tests of a partial airplane. This method requires that all control system components that are attached to or enclosed by the loaded test structure be installed per type design. A sufficiently representative mockup of remaining control system components must be used to ensure that the full length of any cables which extend from the loaded test structure are included. This is necessary to make a reasonable assessment that slack that could develop in control cables is not excessive enough to cause an entanglement or jam. The control surface activation may be input at any convenient location between the mockup terminus and the cockpit.
SECTION (3):
(3) The tests described in this section will demonstrate that the control system is free from excessive deflection as required by § 23.683, paragraph (a)(3). These tests complete this means of compliance that the control system is free from jamming and excessive friction as required by § 23.683, paragraphs (a)(1) and (2). They also demonstrate that structural deformations do not interfere with safe operation as required by § 23.305, paragraph (a). These tests meet the limit load static test requirements of § 23.681, paragraph (a). Accomplish the following:
(i) With the adjacent fixed surface (wing, horizontal tail, or vertical tail) unloaded, support the control surface being tested while it is located in the neutral position.
(ii) Load the control surfaces to the critical limit loads, as described in paragraph f above, and evaluate their proximity to the fixed adjacent structure for jamming or contact.
(iii) Load the pilot's control until the control surface is just off the support.
(iv) Operate the cockpit control in the direction opposite the load to the extent of its travel.
(v) The above procedure should be repeated in the opposite direction.
(vi) The minimum loaded control surface travel must have adequate flight characteristics as specified in § 23.141.
(vii) Under limit load, no signs of jamming, or of any permanent set of any connection, bracket, attachment, and so forth, may be present.
(viii) The control system should operate freely without excessive friction.
NOTE 2: The tests described in section (3) above are normally accomplished using a complete airplane. As a minimum, they must be completed using an airframe/control system that completely represents the final product from the cockpit controls to the control surface.
Regardless of the amount of travel of a control surface when tested as described above, the airplane must have adequate flight characteristics as specified in § 23.141. Any airplane that is a close derivative of a previous type certificated airplane need not exceed the control surface travel of the original airplane; however, the flight characteristics should be tested to ensure compliance.
JAA ACJ 23.683 is acceptable for FAA certification.
23.685 Control system details No policy available as of June 30, 2001.
23.687 Spring devices Compliance may be shown by flight tests with the spring disconnected to demonstrate that failure of the spring will not cause flutter or unsafe flight characteristics, or by performing a reliability analysis.
The flight test option is preferred since this approach conclusively addresses the safety issue and is historically a minimal test burden.
An applicant who chooses the reliability analysis option must accept that a failure of the spring could create unsafe flight characteristics and, therefore, meet the corresponding level of reliability. This may involve the use of redundant design such as dual springs and demonstration of flight characteristics with one removed.
A positive determination of spring reliability requires that an applicant show the spring will perform its intended function for a specified interval under operational and environmental conditions appropriate for the proposed airplane. Although not normally used for structural substantiation, the reliability assessment methods for § 23.1309 may contain some concepts that would be helpful in demonstrating the reliability of the spring device. The reliability assessment should consider, but not necessarily be limited to, fatigue failures, failures due to corrosive environments, and any in-service changes in the spring characteristics, particularly the spring constant. The reliability assessment of a spring device used in any airplane flight control system must consider airworthiness standards other than § 23.687. These include, but are not limited to, flutter characteristics and handling qualities.
In addition, § 23.687 is explicit in that the applicant must demonstrate reliability of the spring with tests that simulate service conditions. Tests are the required substantiation method;
however, an applicant may show compliance with tests supported by analysis. In addition to the spring device testing requirement, 14 CFR part 23, § 23.601, General, requires testing for any design detail or part that has an important bearing on safety of flight. An applicant should also show compliance with 14 CFR part 23, § 23.305 (a) and (b). Finally, when a single spring is required for flutter, handling qualities, or any other regulatory reason, the spring should be considered a single path critical structure and meet the A-Basis requirements of § 23.613(b)(1).
23.689 Cable systems Original Issue and Subsequent If tabs are installed with cable less than 1/8th inch diameter, the airplane should be safely controllable with the tabs in the most adverse position as if from a failed cable. Using emergency procedures, the pilot should be able to return and land safely. Airplane configurations, such as flaps, landing gear, and power are permissible devices to use in relieving control forces. The temporary control forces of part 23, § 23.143, are applicable until the force reduction procedures are completed.
Smaller diameter cables (no less than 1/16 inch diameter) may be used for rudder pedal interconnections (used for pulling one pedal back when the other is pushed forward, but not used to drive the control surface), if the failure of this interconnection will not affect rudder operation.
This rule was put in the original issue of part 23 from CAR 3.345, and it was intended to apply to airframe control cables not engine controls cables, which are certified under part 23, Subpart E.
23.691 Artificial stall barrier system Amendment 23-49 and Subsequent Section 23.201(b), Amendment 23-45, added the activation of an artificial stall barrier as an acceptable means of identifying when a stall has occurred. A stall barrier is a device that prevents an actual stall (i.e., a stick pusher) while a stall warning is a device that alerts a pilot of an impending stall (i.e., a stick shaker). Of course, the actual stall should not occur before activation of the stall barrier. This amendment provided the standards for an artificial stall barrier system when it is used to show compliance with § 23.201(b).
Per § 23.201, a stick shaker is a "stall avoidance" device. Per the NPRM for Amendment 23-45, an applicant may identify a stall as the speed at which a stick shaker activates. This is an acceptable means of compliance to the rules. Stick shaker activation is then identified as a stall for a pilot by AFM. However, from an engineering perspective and considering the "stall avoidance" in § 23.201, this is an acceptable means of compliance not an aerodynamic stall. The difference is important in certification since the aerodynamic stall determines whether a stick shaker is required equipment.
A stick pusher system would be a critical system for an airplane with stall recovery that is undetermined, marginal, or unacceptable. Failure of the system is then required to be extremely improbable. The FAA does not consider the probability of entering a stall environment as a factor in developing system reliability. The exception would be developing specific system component reliability where that component would be active only when the airplane is in a stall environment. The FAA does not give credit toward developing reliability for the use of a ‘Go/No Go’ preflight system check, although the FAA does recommend that preflight procedures for all essential/critical systems be provided for pilot use. (Service experience has shown that some part 23 airplane pilots do not have the discipline to conduct the prescribed preflight checks.) The development of normal/abnormal/emergency procedures is not a factor in determining system reliability;
however, such procedures are desirable, as well as required by § 23.1581. These factors may be considered when exercising engineering judgment in approval of the overall system.
Stall may be identified by stick shaker/pusher operation, uncontrollable downward pitching, or the elevator control reaching the stop (see AC 23-8A)—whichever occurs first in any particular flight regime is acceptable. An airplane may be approved if it has stick shaker/pusher operation in one configuration, such as power on, and it has acceptable stall characteristics for the remaining configurations.
Inadvertent stick pusher operation should be investigated and shown not to be hazardous and to be recoverable, or that inadvertent operation is extremely improbable.
23.693 Joints No policy available as of June 30, 2001.
23.697 Wing flap controls No policy available as of June 30, 2001.
23.699 Wing flap position indicator No policy available as of June 30, 2001.
23.701 Flap interconnection Original Issue The flaps should be synchronized by a mechanical interconnection unless the airplane has safe flight characteristics with the flaps retracted on one side and fully extended on the other side. The safe flight demonstration with asymmetry should be shown throughout the airspeed range permitted for flap extension. The control forces should not exceed those shown for temporary application in the table in § 23.143(c). However, they may not exceed the force that can be demonstrated as safe with one hand on the control wheel/stick (other hand needed to re-trim, pull circuit breaker, operate flap control, etc.). If the forces of asymmetry cannot be alleviated in a reasonable period of time, the remaining forces should not exceed those specified for prolonged application in § 23.143(c).
After demonstrating that the airplane has safe flight characteristics with the flaps in their most adverse position, it is permissible to readjust the remaining flap surfaces after a malfunction occurs.
Amendment 23-42 Amendment 23-42 was not intended to change the requirement that “The main wing flaps and related movable surfaces as a system must be synchronized by mechanical connection.
. . .” The main purpose of this change was to add the following requirement that would maintain synchronization so that the occurrence of an unsafe condition has been shown to be extremely improbable. This requirement includes provisions for synchronization of the flaps other than by mechanical interconnection of the flap. These reliability requirements by numerical probability analysis for other synchronization methods should not be applied to mechanical interconnection.
It is difficult to assess the reliability of mechanical interconnections by examples of different types of mechanisms. The complete system needs to be analyzed and tested.
Section 23.701, as amended by Amendment 23-42, in part, states the following:
(a) The main wing flaps and related movable surfaces as a system must:
(1) Be synchronized by mechanical connection; or
(2) Maintain synchronization so the occurrence of an unsafe condition has been shown to be extremely improbable; or