Cargas devidas a Manobras de Guinada e Rajadas Laterais

1. ITA – Instituto Tecnológico de Aeronáutica Cargas devidas a Manobras de Guinada e Rajadas Laterais Cargas em Aviões

2. Tipos de Manobras de Guinada. As manobras de leme usadas para o projeto estrutural são essencialmente manobras “rasas”, onde o leme é aplicado bruscamente, numa atitude de asas niveladas. Esta manobra é difícil de realizar em vôo porque controle lateral considerável deve ser aplicado para manter as asas niveladas. O propósito de se manter as asas niveladas é maximizar a derrapagem. As manobras com falha de motor, como usadas para o projeto estrutural, são manobras essencialmente “rasas”, onde uma aplicação abrupta do leme é feita em conjunção com a derrapagem resultante da tração assimétrica do motor.

3. Part 25 § 25.351 - Yaw maneuver conditions. 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 VMC to VD. 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 or control surface stops; or (2). A limit pilot force of 300 pounds from VMC to VA and 200 pounds from VC/MC to VD/MD, with a linear variation between VA and VC/MC. (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.

4. Part 25 § 25.149 – Minimum control speed. (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). VMC 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). VMC may not exceed 1.2 Vs 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 VMC); (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 VMC 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.

5. Part 25 § 25.351 - Yaw maneuver conditions. Condição não simétrica Fator de carga simétrico Condição assimétrica manobras de guinada 1 Velocidades e acelerações de guinada; ângulos de leme e de derrapagem Parâmetros requeridos para análise não-simétrica

6. Part 25 § 25.351 - Yaw maneuver conditions. Man II Man I Man III rudder angle rudder and sideslip angle steady sideslip sideslip angle time Lvr V Maneuver I dr I b≈ 0 V Lvb Lvr b = bmax Maneuver II dr I V Lvb dr = 0 b = bss Maneuver III

7. Leme disponível. A máxima deflexão do leme disponível a uma dada velocidade e número de Mach é função da força do piloto e da unidade de controle de potência. Exemplos do comportamento de unidades de controle de potência são:

8. Steady sideslip due to rudder. dr known sideslip, rudder and aileron control wheel angles side force coefficient due to sideslip, rudder and aileron yawing moment coefficient due to sideslip, rudder and aileron rolling moment coefficient due to sideslip, rudder and aileron lift coefficient bank angle Assuming level flight at constant airspeed and aerodynamic coefficients varying lineraly, the equations for side force, yawing and rolling moments may be written as:

9. Part 25 § 25.367 - Unsymmetrical loads due to engine failure. (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 VMC and VD, the loads resulting from power failure because of fuel flow interruption are considered to be limit loads. (2). At speeds between VMC and Vc, 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 enginepropeller 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 analysis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions.

10. Part 25 § 25.397 – Control system loads. (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 Footnote: 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. Footnote: 2 - D = wheel diameter (inches). Footnote: 3 - The unsymmetrical forces must be applied at one of the normal handgrip points on the periphery of the control wheel.

11. Part 25 § 25.367 - Unsymmetrical loads due to engine failure. • Para cada velocidade em consideração, devem ser examinadas as condições: • Derrapagem máxima produzida com zero leme; e • Leme corretivo aplicado não antes de 2 segundos ou no momento da máxima velocidade de guinada. • Condições de projeto: • a falha de motor devida à interrupção de alimentação de combustível deve ser considerada na condição de carga limite e, portanto, o fator de segurança 1,5 deve ser aplicado; • a falha de motor devida a problemas mecânicos dos sistemas do motor ou da hélice deve ser considerada como uma condição de carga final, de modo que o fator de segurança a ser aplicado é 1,0. A diferença entre estas duas condições é o tempo de decaimento da tração do motor. A interrupção de combustível pode ocorrer desde um, a vários segundos, enquanto que a falha mecânica ocorre abruptamente.

12. Engine-out analysis steady-state conditions Engine-out steady sideslip with zero rudder (dr = 0) Engine-out steady condition with zero sideslip (beo = 0)

13. Equations of motion for yawing maneuvers (3 DOF model) • Since the yawing maneuvers for structural load analyses are considered “flat”: • Roll accleration and velocity are assumed zero; • Lateral control is applied as necessary to maintain a wings-level atitude; • Airspeed and Mach number (hence altitude) are assumed constant during the maneuver; • Rate derivatives of the rudder and lateral control devices are neglected. Dividing the side force equation by qSw and the roll and yaw moment equations by qSwbw, one obtains

14. Equations of motion for yawing maneuvers (2 DOF model) Neglecting the roll degree of freedom results in

15. Rudder maneuver analysis Numerical integration solution of 3-DOF and 2-DOF set of equations, neglecting de engine-out term. Ramp rudder input motion, with time for 7.2 degree maximum rudder reached at 0,18 s. Analyses show abrupt rudder condition, maneuver I, and the maximum overway condition, maneuver II.

16. Engine-out maneuver analysis The engine-out analysis is usually done in two steps: 1- using thrust decay as input, and assuming no corrective action by the pilot with the rudder, one runs the time history analysis to determine the maximum sideslip angle of the airplane and the time for maximum yaw rate; 2- the analysis is now rerun with corrective rudder initiated at the time of maximum yaw rate (no sooner than 2 seconds)

17. Engine-out maneuver analysis (3 DOF model) Engine-out maneuver analysis with no corrective rudder action.

18. Engine-out maneuver analysis (3 DOF model) Engine-out maneuver analysis with corrective rudder action initiated at 2.0 sec; the corrective rudder is as required for zero sideslip in the steady-state condition.

19. Rajadas (FAR Part 25 §25.341) (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 following 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) s = distance penetrated into the gust (feet); Ude = the design gust velocity in equivalent airspeed specified in paragraph (a)(4) of this section; 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 Ude = Uref x Fg(H/350)1/6 , where Uref = the reference gust velocity in equivalent airspeed defined in paragraph (a)(5) of this section. Fg = the flight profile alleviation factor defined in paragraph (a)(6) of this section.

20. Rajadas (FAR Part 25 §25.341) (5). The following reference gust velocities apply: (i). At the airplane design speed VC: 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 15000 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at 15000 feet to 26.0 ft/sec EAS at 50000 feet. (ii). At the airplane design speed VD: The reference gust velocity must be 0.5 times the value obtained under § 25.341(a)(5)(i) (6). The flight profile alleviation factor, Fg, 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: Fg = 0.5(Fgz + Fgm) , where Zmo = maximum operating altitude defined in § 25.1527. (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.

21. Rajadas (FAR Part 25 §25.341 and §25.427) (b). Continuous Gust Design Criteria. The dynamic response of the airplane to vertical and lateral continuous turbulence must be taken into account. The continuous gust design criteria of appendix G of this part must be used to establish the dynamic response unless more rational criteria are shown. § 25.427 -- Unsymmetrical loads. (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 § 24.305(e) must be taken into account.

22. Fórmula da Carga devida à Rajada Lateral A carga devida à rajada lateral, assumindo um corpo rígido, pode ser representada por onde Kg é o fator de alívio da rajada com mv, a razão de massa lateral, dada por velocidade derivada de rajada (ft/s eas) velocidade equivalente do avião (nós eas) inclinação da curva de sustentação da empenagem vertical (rad-1)s superfície de referência da empenagem vertical (ft2) momento de inércia em arfagem (slug ft2) densidade do ar no nível do mar (slug/ft3) corda média aerodinâmica da empenagem vertical (ft) distância do CA da empenagem vertical ao CG (ft)

23. Componentes de Carga devidas à Rajada Oblíqua Ude Udesenf Udecosf componente lateral da carga nas empenagens devida à rajada oblíqua componente vertical da carga nas empenagens devida à rajada oblíqua carga nas empenagens devida ao vôo 1-g carga de projeto nas empenagens devida à rajada lateral de intensidade Ude carga de projeto nas empenagens devida à rajada vertical de intensidade Ude