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Cargas devidas a Manobras e Rajadas Simétricas Cargas em Aviões ITA – Instituto Tecnológico de Aeronáutica.

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Apresentação em tema: "Cargas devidas a Manobras e Rajadas Simétricas Cargas em Aviões ITA – Instituto Tecnológico de Aeronáutica."— Transcrição da apresentação:

1 Cargas devidas a Manobras e Rajadas Simétricas Cargas em Aviões ITA – Instituto Tecnológico de Aeronáutica

2 Cargas devidas a manobras e rajadas simétricas Rajadas Simétricas Manobras Simétricas Manobra Diagrama V-n de Manobras Velocidades de Projeto Fatores de Carga Típicos Carga de Balanceamento na Empenagem Horizontal Rajadas de contorno vivo Rajadas com gradiente FAR Part 23 § Diagrama V-n do Manobras e Rajadas FAR Part 25 § Efeitos da rajada na empenagem horizontal Efeitos Estruturais das Manobras e Rajadas Simétricas

3 Manobras Simétricas Linha da corda da asa Eixo x do avião Horizontal Trajetória

4 Nomenclatura sustentação do avião, excluída a contribuição da cauda arrasto do avião, excluída a contribuição da cauda momento aerodinâmico de avião, excluída a contribuição da cauda, em torno do CA sustentação da cauda arrasto da cauda momento aerodinâmico da cauda em torno do CA da cauda força aerodinâmica do avião, excluída a contribuição da cauda, na direção z força aerodinâmica do avião, excluída a contribuição da cauda, na direção x força aerodinâmica da cauda na direção z força aerodinâmica da cauda na direção x força de tração do motor coordenada z do centro aerodinâmico do avião, excluída a contribuição da cauda coordenada x do centro aerodinâmico do avião, excluída a contribuição da cauda coordenada z do CA da cauda coordenada x do CA da cauda coordenada z do motor coordenada x do motor aceleração do avião na direção z aceleração do avião na direção x aceleração angular do avião em arfagem

5 aceleração da gravidade superfície da asa pressão dinâmica corda média aerodinâmica coeficiente de sustentação do avião, excluída a contribuição da cauda coeficiente de arrasto do avião, excluída a contribuição da cauda fator de carga na direção z fator de carga na direção x ângulo de ataque (entre a trajetória e a linha de referência da corda média) ângulo entre a trajetória e o eixo x do avião ângulo entre a o eixo x e a linha de ação da tração do motor ângulo de arfagem (entre a horizontal e o eixo x do avião) ângulo entre a horizontal e a trajetória ângulo de incidência da asa (entre o eixo x e linha de referência da corda média) ângulo de incidência da empenagem horizontal Nomenclatura Decomposição das forças aerodinâmicas totais do avião nos eixos z e x

6 Manobras Simétricas - Simplificações Pelo menos na fase de anteprojeto, o momento aerodinâmico da cauda, bem como a contribuição de F xT, podem ser desprezados na equação do momento. Por outro lado, normalmente podem também ser desprezados os ângulos i w e  F. Equilíbrio do avião em trajetória horizontal uniforme, com aceleração de arfagem nula. (condição de aceleração nula na horizontal)

7 Manobras Simétricas - Simplificações Equilíbrio do avião em trajetória horizontal uniforme, com aceleração de arfagem nula. Dados n z, W, S, q, os parâmetros geométricos, e conhecida a variação dos coeficientes aerodinâmicos com , pode-se determinar o ângulo de ataque  (e.g., usando o MatLab ou Excel). Os demais parâmetros podem, então, ser calculados:  C C za C zw C xa C zT C MAC

8 Exemplo 2 Uma aeronave com flapes na posição neutra tem as seguintes características: a) Asa retangular  CLCL CDCD C mc/4 13,6 o 1,340,161-0,09 8o8o 0,970,062-0,09 4o4o 0,660,032-0,09 0o0o 0,320,014-0,09 -4 o 0,020,011-0,09 -8 o -0,30,022-0, o -0,620,041-0, o -0,960,098-0,09 -1,5 -0,5 0 0,5 1 1,  CLCL CDCD C mc/4

9 b) Coeficiente de arrasto do avião, excetuando a asa: c) Momento de arfagem da fuselagem, excetuando a asa: d) Posição do CA da asa em relação ao CG: e) Posição do CA da empenagem em relação ao CG: f) Posição do motor em relação ao CG: g) Peso da aeronave: 1. Ache os coeficientes C zw, C xw, C zT, C za, C xa e C m em função de , na condição de tração nula 2. Ache as forças F zw, F xw, F zT, F za, F xa e F eng, e o fator de carga n x, na condição de vôo horizontal não acelerado, com n z = 5,7 e V E = 250 m/s (  0 = 1,223 kg/m 3 ) Exemplo 2

10 Na condição de tração e aceleração de arfagem nulas, a equação de equilíbrio dos momentos fica Desprezando as contribuições de M T e F xT, e levando em consideração que no enunciado M AC é dado em duas parcelas: Dividindo a equação por Sq resulta em

11 Exemplo 2  CLCL CDCD C mc/4 C mf C xw C zw C zT C za C xa CmCm 13,61,340,161-0,09 0,028-0,1591,340-0,0981,242-0,126-0,130 80,970,062-0,09 0,016-0,0740,969-0,0920,877-0,041-0,058 40,660,032-0,09 0,007-0,0140,661-0,0870,5740,019-0,007 00,320,014-0,09 -0,0020,0140,320-0,0790,2410,0470, ,020,011-0,09 -0,0100,0120,019-0,071-0,0520,0450, ,30,022-0,09 -0,019-0,020-0,300-0,061-0,3610,013-0, ,620,041-0,09 -0,028-0,089-0,615-0,049-0,664-0,056-0, ,960,098-0,09 -0,039-0,187-0,947-0,036-0,982-0,154-0,226

12 Exemplo 2

13  (graus) CLCL CDCD C mc/4 C mf CmCm mão esq.mão dir. 13,61,340,161-0,09 0,028-0, ,970,062-0,09 0,016-0, ,660,032-0,09 0,007-0, ,320,014-0,09 -0,002-0, ,020,011-0,09 -0,010-0, ,30,022-0,09 -0,019-0, ,620,041-0,09 -0,028-0, ,960,098-0,09 -0,039-0, Calculando-se a mão esquerda da equação acima em função de , com C MAC = C MAC/4 + C Mf, obtém- se Como pode ser notado, o ângulo solução está entre -4 o e 0 o.

14 Exemplo 2  (graus) CLCL CDCD C mc/4 C mf CmCm mão esq.mão dir. -2, , , ,09-0,008-0, ,00 F zw (N)F xw (N)F xa (N)F eng (N)F zt (N)F za (N)nxnx ,285 Efetuando uma interpolação quadrática para C L e C D, usando os dados para  = 0 o, -4 o e -8 o, obtém-se Resolvendo a equação de equilíbrio de momentos, obtém-se  = -2, o

15 Manobras Simétricas Em vôo nivelado não-acelerado, o piloto puxa o manche maior ângulo de ataque maior fator de carga vertical Hipótese: Velocidade constante.

16 Diagrama V-N de Manobras Envoltória das condições permitidas para operação normal do avião. velocidade verdadeira nível do mar velocidade equivalente

17 Diagrama V-N de Manobras Flight maneuvering envelope (FAR Part 25§ )

18 Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ ) The selected design airspeeds are equivalent airspeeds (EAS). V S0 = power off stalling speed at sea level, at the design landing weight, and in the landing configuration; V S1 = stalling speed with flaps retracted; V A = design maneuvering speed; V B = design speed for maximum gust intensity; V C = design cruising speed; V D = design diving speed; V F = design flap speeds; V DD = design drag device speed.

19 Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ ) ( a). 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 § (d)(2), V C may not be less than V B U REF (with U REF as specified in § (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. § (a)(5)(i). At the airplane design speed V C, positive and negative gusts with reference gust velocities (U REF ) 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 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at feet to 26.0 ft/sec EAS at feet.

20 Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ ) ( b). 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 7.50 degrees below the initial path, and then pulled up at a load factor of 1.5g (0.5g acceleration increment). The speed increase occurring in this maneuver may be calculated if reliable or conservative aerodynamic data is issued. Power as specified in § (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 be 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.

21 (c). For V A, the following apply: (1). V A may not be less than where n is the limit positive maneuvering load factor at V C ; and V S1 is the stalling speed with flaps retracted. (2). V A and V S1 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 Nmax curve intersects the positive maneuver load factor line, whichever is less. Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ )

22 ( d). (1). V B may not be less than where: K g = gust alleviation factor = m g = airplane mass ratio = 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 § (a)(5)(i); w = average wing loading (pounds per square foot) at the particular weight under consideration.  = density of air (slug/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. Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ )

23 ( e). For V F, the following apply: (1). The design flap speed for each flap position (established in accordance with § (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 S1, with the flaps in takeoff position at maximum takeoff weight; (ii). 1.8 V S1, with the flaps in approach position at maximum landing weight; and (iii). 1.8 V S0 with the flaps in landing position at maximum landing weight. (f). The selected design speed for each drag device (V DD ) 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. Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ )

24 (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 V D may not be less than 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 in pounds (2.5 for W > 60,000 lb ; 3.8 for W < 4,118 lb) (c). The negative limit maneuvering load factor- (1). May not be less than 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. Diagrama V-N de Manobras Limit maneuvering load factors (FAR Part 25§ )

25 Fatores de Carga Típicos

26 Part 25 § Symmetric maneuvering conditions (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 § 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 § (b) must be investigated. (c). Pitch maneuver conditions. The conditions specified in paragraphs (c)(1) and (2) of this section must be investigated. The movement of the pitch control surfaces may be adjusted to take into account limitations imposed by the maximum pilot effort specified by § (b), control system stops and any indirect effect imposed by limitations in the output side of the control system (for example, stalling torque or maximum rate obtainable by a power control system.) (1). Maximum pitch control displacement at V A. The airplane is assumed to be flying in steady level flight (point A1,§ (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 A2 in § (b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.

27 Part 25 § Symmetric maneuvering conditions ( 2). Specified control displacement. A checked maneuver, based on a rational pitching control motion vs. time profile, must be established in which the design limit load factor specified in § will not be exceeded. Unless lesser values cannot be exceeded, the airplane response must result in pitching accelerations not less than the following: (i). A positive pitching acceleration (nose up) is assumed to be reached concurrently with the airplane load factor of 1.0 (points A1 to D1, § (b)). The positive acceleration must be equal to at least 39 n ( n – 1.5 ) / V Radians/sec2 Where n is the positive load factor at the speed under consideration; and V is the airplane equivalent speed in knots. (ii). A negative pitching acceleration (nose down) is assumed to be reached concurrently with the positive maneuvering load factor (points A2 to D2, §25.333(b)). This negative pitching acceleration must be equal to at least -26 n ( n – 1.5 ) / V Radians/sec2

28 Carga de Balanceamento da Empenagem Horizontal (F zT ) B é a carga na empenagem horizontal necessária para manter o avião em equilíbrio, com aceleração de arfagem nula Na condição sem motor, desprezando como pequenas as parcelas M T e F xT (z T -z w ), obtém-se

29 Carga de Balanceamento da Empenagem Horizontal nznz VEVE A D E H (F zT ) B H E D VEVE A VEVE A D E H

30 Distribuição de (F zT ) B na Empenagem Horizontal pp TT estabilizador profundor Distribuição típica de pressão  p = 10 o  T = 0 o  p = 0 o  T = 1 0 o Características aerodinâmicas da empenagem  p = 0 o 10 o 20 o -10 o -20 o TT C´ zT pp

31 Manobras não Corrigidas (Unchecked Maneuvers) nznz VEVE A D E H n z =1 A1A1 A2A2 VAVA (1). Maximum pitch control displacement at V A. The airplane is assumed to be flying in steady level flight (point A1,§ (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 A2 in § (b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.

32 Manobras Corrigidas (Checked Maneuvers) (2). Specified control displacement. A checked maneuver, based on a rational pitching control motion vs. time profile, must be established in which the design limit load factor specified in § will not be exceeded. Unless lesser values cannot be exceeded, the airplane response must result in pitching accelerations not less than the following: (i). A positive pitching acceleration (nose up) is assumed to be reached concurrently with the airplane load factor of 1.0 (points A1 to D1, § (b)). The positive acceleration must be equal to at least 39 n ( n – 1.5 ) / V Radians/sec 2 Where n is the positive load factor at the speed under consideration; and V is the airplane equivalent speed in knots. (ii). A negative pitching acceleration (nose down) is assumed to be reached concurrently with the positive maneuvering load factor (points A2 to D2, §25.333(b)). This negative pitching acceleration must be equal to at least -26 n ( n – 1.5 ) / V Radians/sec 2 nznz VEVE A D E H n z =1 A1A1 A2A2 VAVA D2D2 D1D1

33 Manobras não Corrigidas – Análise Aproximada nznz VEVE A D E H n z =1 A1A1 A2A2 VAVA Antes da manobra Depois de  p  p = 0 o 10 o 20 o -10 o -20 o TT C´ zT A1A1 A2A2

34 Manobras Corrigidas – Análise Aproximada nznz VEVE A D E H n z =1 A1A1 A2A2 VAVA D2D2 D1D1 Carga para baixo [n z = 1;  = 39 n max (n max - 1) / V ; V dado em knots ] Carga para cima [n z = n max ;  = -26 n max (n max - 1) / V - V dado em knots ]

35 Rajadas Rajadas de contorno vivo Rajadas com gradiente FAR Part 23 § Diagrama V-n de Manobras e Rajadas FAR Part 25 § Efeitos de rajadas na empenagem horizontal

36 Rajadas de Contorno Vivo

37 Variação instantânea: VEVE UEUE VRVR 

38 a- o deslocamento vertical do avião reduz o efeito da rajada e a rotação em arfagem; b- efeito de Wagner: retardo na variação de C zw com  ; c- efeitos dinâmicos (aeroelásticos) no avião. Cálculo aproximado: considerar a rajada com gradiente, como uma rajada de contorno vivo à qual se associa um fator de alívio K g Rajadas com Gradiente (Graded Gust)

39 Rajadas (FAR Part 23 §23.341) U de = Derived gust velocities referred to in § (c)(f.p.s.);  0 = Density of air (slugs/cu.ft.); W/S = Wing loading (p/sq.ft.) due to the applicable weight of the airplane in the particular load case. c = Mean geometric chord (ft.); g = Acceleration due to gravity (ft./Sec. 2 ) V = Airplane equivalent speed (knots); and a = Slope of the airplane normal force coefficient curve C NA per radian if the gust loads are applied to the wings and horizontal tail surfaces simultaneously by a rational method. The wing lift curve slope C L per radian may be used when the gust load is applied to the wings only and the horizontal tail gust loads are treated as a separate condition. § (c). Gust envelope. (1). The airplane is assumed to be subjected to symmetrical vertical gusts in level flight. The resulting limit load factors must correspond to the conditions determined as follows: (i). Positive (up) and negative (down) gusts of 50 f.p.s. at V C must be considered at altitudes between sea level and 20, 000 feet. The gust velocity may be reduced linearly from 50 f.p.s. at 20, 000 feet to 25 f.p.s. at 50, 000 feet. (ii). Positive and negative gusts of 25 f.p.s. at V D must be considered at altitudes between sea level and 20, 000 feet. The gust velocity may be reduced linearly from 25 f.p.s. at 20,000 feet to 12.5 f.p.s. at 50, 000 feet. (iii). In addition, for commuter category airplanes, positive (up) and negative (down) rough air gusts of 66 f.p.s. at V B must be considered at altitudes between sea level and 20, 000 feet. The gust velocity may be reduced linearly from 66 f.p.s. at 20, 000 feet to 38 f.p.s. at 50, 000 feet.

40 Diagrama V-n de Manobras e Rajadas FAR Part 23 § Válido para um dado peso (W) e posição do CG do avião

41 (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) Rajadas (FAR Part 25 §25.341) s = distance penetrated into the gust (feet); U de = 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.

42 Rajadas (FAR Part 25 §25.341) ( 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 U de = U ref F g (H/350) 1/6, 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 the airplane design speed 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 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at feet to 26.0 ft/sec EAS at feet. (ii). At the airplane design speed V D : The reference gust velocity must be 0.5 times the value obtained under § (a)(5)(i)

43 Rajadas (FAR Part 25 §25.341) (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 § At sea level, the flight profile alleviation factor is determined by the following equation: Fg = 0.5(F gz + F gm ), where Z mo = maximum operating altitude defined in § (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 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.

44 Efeitos de Rajadas na Empenagem Horizontal C´ zT TT  T  C´ zT Na asa: Na empenagem: No avião:

45 FAR Part 25 § – Design fuel and oil loads. (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 § (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 conditions of § (a) but assuming 85% of the design velocities prescribed in § (a)(4). (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. Note: § (f) states that there should be enough fuel in the tanks used for takeoff and landing, to allow climb from sea level to 10,000 feet and 45 minutes cruise at a speed for maximum range thereafter.

46 FAR Part 25 § – High lift devices. (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 § (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 § (a)(2) except that- Uds = 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 VF, 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 discrete vertical gust criteria in § (a). (d). The airplane must be designed for a maneuvering load factor of 1.5 g at the maximum takeoff weight with the wing-flaps and similar high lift devices in the landing configurations.

47 Efeitos Estruturais das Manobras Simétricas Ponto A: alto ângulo de ataque positivo nznz VEVE A D E H L  D F zw F xw Ponto D: baixo ângulo de ataque positivo L  D F zw F xw (-) (-) (-) (+) (+) (+) (-) (-) (-) (+) (+) (+) M Fzw M Fxw críticos M Fzw (-) (-) (-) (-) (-) (-) (+) (+) (+) (+) (+) (+) M Fxw críticos

48 Efeitos Estruturais das Manobras Simétricas nznz VEVE A D E H Torção Ao longo de AD variam F zw e F xw, embora F za = n z W = cte. As forças de inércia, que dependem de n z, são constantes, de modo que a flexão na asa varia ao longo de AD. y yy  M y máximo positivo para V E mínimo e n z máximo positivo → Ponto A do diagrama V-n  M y máximo negativo para V E máximo e n z máximo negativo → Ponto E do diagrama V-n CA CG CT  F zw  F xw  M CA a b nxWnxW nzWnzW MyMy Flexão nas asas

49 Variação das Cargas com o Peso e Posição do CG CG CA CA T nzWnzW F zw M CA F zT xwxw xTxT nzWfnzWf Ex. Esforços Máximos em Seção da Fuselagem Traseira

50 Rajada com Peso Mínimo F zw é máxima para CG à frente e W max n z é mínimo para W max Flexão é máxima Para itens de equipamento cujas cargas dependem fundamentalmente das forças de inércia, a condição de rajada com peso mínimo poderá ser crítica (e.g., berços de motores)

51 Carga Variável na Asa Flexão pode ser crítica com peso abaixo do máximo a) Com combustível M i = contribuição dos demais itens de inércia da asa b) Sem combustível combustível l1l1 l2l2 n z W comb /2 n z W/2 A A Peso do avião = W - W comb


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