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

Cargas devidas a Manobrase Rajadas Simétricas

Cargas em Aviões

ITA – Instituto Tecnológico de Aeronáutica

Cargas devidas a manobras e rajadas simétricas

Rajadas Simétricas

Manobras SimétricasManobraDiagrama V-n de ManobrasVelocidades de ProjetoFatores de Carga TípicosCarga de Balanceamento na Empenagem Horizontal

Rajadas de contorno vivoRajadas com gradienteFAR Part 23 §23.341Diagrama V-n do Manobras e RajadasFAR Part 25 §25.341Efeitos da rajada na empenagem horizontal

Efeitos Estruturais das Manobras e Rajadas Simétricas

Manobras Simétricas

qcSCM

SqCD

SqCL

DLF

LDF

MACAC

D

L

wz

wx

sencos

sencos

Linha da corda da asa

Eixo x do avião

Horizontal

Trajetória

wi

; ;

MACMACDDLL

w

CCCCCC

i

wz

wx

ti

wi

Fz

Fx

x

z

zF wzF

xF wxF

L

D

ACMTM

TzF

TxF

Vento

Tx

Tz

za

xaengF

Wza

g

W

xag

W

yI

TL

TD

F

yTTTxTTzFFengFFengACwwxwwz

zzFengazzFengTzwz

xxFengaxxFengTxwx

IMzFxFxFzFMzFxF

WngaWFFagWWFFF

WngaWFFagWWFFF

sencos

cossen cossen

sencossencos

Nomenclaturasustentação do avião, excluída a contribuição da caudaarrasto do avião, excluída a contribuição da caudamomento aerodinâmico de avião, excluída a contribuição da cauda, em torno do CAsustentação da caudaarrasto da caudamomento aerodinâmico da cauda em torno do CA da caudaforça aerodinâmica do avião, excluída a contribuição da cauda, na direção zforça aerodinâmica do avião, excluída a contribuição da cauda, na direção xforça aerodinâmica da cauda na direção zforça aerodinâmica da cauda na direção xforça de tração do motorcoordenada z do centro aerodinâmico do avião, excluída a contribuição da caudacoordenada x do centro aerodinâmico do avião, excluída a contribuição da caudacoordenada z do CA da caudacoordenada x do CA da caudacoordenada z do motorcoordenada x do motoraceleração do avião na direção zaceleração do avião na direção xaceleração angular do avião em arfagem

x

z

F

F

T

T

w

w

eng

Tx

Tz

wx

wz

T

T

T

AC

a

a

x

z

x

z

x

z

F

F

F

F

F

M

D

L

M

D

L

aceleração da gravidadesuperfície da asapressão dinâmicacorda média aerodinâmicacoeficiente de sustentação do avião, excluída a contribuição da caudacoeficiente de arrasto do avião, excluída a contribuição da caudafator de carga na direção zfator de carga na direção xângulo de ataque (entre a trajetória e a linha de referência da corda média)ângulo entre a trajetória e o eixo x do aviãoângulo entre a o eixo x e a linha de ação da tração do motorângulo de arfagem (entre a horizontal e o eixo x do avião)ângulo entre a horizontal e a trajetóriaângulo de incidência da asa (entre o eixo x e linha de referência da corda média)ângulo de incidência da empenagem horizontal

Nomenclatura

T

w

F

x

z

D

L

i

i

n

n

C

C

c

q

S

g

Txwxax

Tzwzaz

FFF

FFF

Decomposição das forças aerodinâmicas totais do avião nos eixos z e x

Manobras Simétricas - Simplificações

T

FengwDLwLDMACTz

DLwz

LDwx

x

zFSqxCCSqzCCSqcCF

SqCCDLF

SqCCLDF

sencossencos

sencossencos

sencossencos

0sencos ;0 ;0 ;0

0 ; 0

T

w

zxF aaαM

i

Wnx

zFSqxxCCzCCcC

nWFSqCCC

zT

FengwTDLwLDMAC

zengTxLD

sencossencos

tancossensencos

0

coscos cos

sensen

TTzFengACwwxwwz

zzzzazzTzwz

xxengaxxengax

xFzFMzFxF

gnaWngaWFagWWFF

WngaWFFagWWFF

Pelo menos na fase de anteprojeto, o momento aerodinâmico da cauda, bem como a contribuição de FxT , podem ser desprezados na equação do momento. Por outro lado, normalmente podem também ser desprezados os ângulos iw e F.

Equilíbrio do avião em trajetória horizontal uniforme, com aceleração de arfagem nula.

(condição de aceleração nula na horizontal)

tancostan zzx ngaa

Manobras Simétricas - Simplificações

Wn

x

zFSqxxCCzCCcC

nWFSqCCC

zT

FengwTDLwLDMAC

zengTxLD

sencossencos

tancossensencos

Equilíbrio do avião em trajetória horizontal uniforme, com aceleração de arfagem nula.

tan

cos

tancossen

sencos

sencos

zx

zz

engwxx

MACAC

D

L

Tzwzaz

T

FengACwwxwwzTz

xazeng

Txwxax

LDwx

DLwz

aa

gna

WFFn

SqcCM

SqCD

SqCL

FFF

x

zFMzFxFF

FnWF

FFF

SqCCF

SqCCF

Dados nz, 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:

CCza Czw

CxaCzT

CMAC

Wn

x

znWSqxxCCzCzzCCcCz

T

FzwTDLFTxFwLDMAC tancossensencossencos

Exemplo 2

wi

Uma aeronave com flapes na posição neutra tem as seguintes características:

a) Asa retangular

1

6SbA

m 295,1

m 755,9

0

2

2

c

S

iw

CL CD Cmc/4

13,6o 1,34 0,161 -0,09

8o 0,97 0,062 -0,09

4o 0,66 0,032 -0,09

0o 0,32 0,014 -0,09

-4o 0,02 0,011 -0,09

-8o -0,3 0,022 -0,09

-12o -0,62 0,041 -0,09

-17o -0,96 0,098 -0,09

-1,5

-1

-0,5

0

0,5

1

1,5

-20 -10 0 10 20

CL

CD

Cmc/4

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 Czw, Cxw, CzT, Cza , Cxa e Cm em função de , na condição de tração nula

2. Ache as forças Fzw, Fxw, FzT, Fza, Fxa e Feng , e o fator de carga nx, na condição de vôo horizontal não acelerado, com nz = 5,7 e VE = 250 m/s (0 = 1,223 kg/m3)

Exemplo 2

m 153,0 ; m 09,0 ww zx

graus) em ( 0022,00015,0 fmC

NW 230.4

033,0 fxTx CC

m 348,3tx

m 028,0 ; m 71,1 FF zx

Exemplo 2

yTTTxTTzFFengFFengACwwxwwz IMzFxFxFzFMzFxF sencos

0 TTTxTTzACwwxwwz MzFxFMzFxF

T

fCAwwxwwz

TzTTzfCAwwxwwz x

MMzFxFFxFMMzFxF

4

4 0

Na condição de tração e aceleração de arfagem nulas, a equação de equilíbrio dos momentos fica

Desprezando as contribuições de MT e FxT, e levando em consideração que no enunciado MAC é dado em duas parcelas:

Dividindo a equação por Sq resulta em

T

MfCMAwwxwwz

Tz x

cCcCzCxCC

4

senCcosCC DLzw

cosCsenCC DLxw

ztzwza CCC

xfxtxwxa CCCC

MfMACm CCC 4/

Exemplo 2

CL CD Cmc/4 Cmf Cxw Czw CzT Cza Cxa Cm

13,6 1,34 0,161 -0,09 0,028 -0,159 1,340 -0,098 1,242 -0,126 -0,130

8 0,97 0,062 -0,09 0,016 -0,074 0,969 -0,092 0,877 -0,041 -0,058

4 0,66 0,032 -0,09 0,007 -0,014 0,661 -0,087 0,574 0,019 -0,007

0 0,32 0,014 -0,09 -0,002 0,014 0,320 -0,079 0,241 0,047 0,013

-4 0,02 0,011 -0,09 -0,010 0,012 0,019 -0,071 -0,052 0,045 0,002

-8 -0,3 0,022 -0,09 -0,019 -0,020 -0,300 -0,061 -0,361 0,013 -0,039

-12 -0,62 0,041 -0,09 -0,028 -0,089 -0,615 -0,049 -0,664 -0,056 -0,117

-17 -0,96 0,098 -0,09 -0,039 -0,187 -0,947 -0,036 -0,982 -0,154 -0,226

Exemplo 2

Coeficientes Aerdinâmicos

-1,5

-1,0

-0,5

0,0

0,5

1,0

1,5

-20 -10 0 10 20

(graus)

C

Cxw

Czw

CzT

Cza

Cxa

Cm

Exemplo 2

(graus) CL CD Cmc/4 Cmf Cm mão esq. mão dir.

13,6 1,34 0,161 -0,09 0,028 -0,062 480.423 24.111

8 0,97 0,062 -0,09 0,016 -0,074 342.319 24.111

4 0,66 0,032 -0,09 0,007 -0,083 227.917 24.111

0 0,32 0,014 -0,09 -0,002 -0,092 102.516 24.111

-4 0,02 0,011 -0,09 -0,010 -0,100 -7.842 24.111

-8 -0,3 0,022 -0,09 -0,019 -0,109 -124.297 24.111

-12 -0,62 0,041 -0,09 -0,028 -0,118 -238.387 24.111

-17 -0,96 0,098 -0,09 -0,039 -0,129 -358.328 24.111

Calculando-se a mão esquerda da equação acima em função de , com CMAC = CMAC/4 + CMf, obtém-se

Como pode ser notado, o ângulo solução está entre -4o e 0o.

Wnx

znWSqxxCCzCCzzCCcC

nWFSqCCCC

Wnx

zFSqxxCCzCCcC

zT

FzwTDLFfxTxFwLDMAC

zengfxTxLD

zT

FengwTDLwLDMAC

tancossensencossencos

tancossensencos

sencossencos

Exemplo 2 400004375,0000075,0014,0

40000625,00075,032,0

xxC

xxC

D

L

(graus) CL CD Cmc/4 Cmf Cm mão esq. mão dir.

-2,865485 0,107121 0,010429 -0,09 -0,008 -0,098 24.111,00 24.111,00

Fzw (N) Fxw (N) Fxa (N) Feng (N) Fzt (N) Fza (N) nx

39.693 5.880 18.183 16.976 -15.582 24.111 0,285

Efetuando uma interpolação quadrática para CL e CD, usando os dados para = 0o, -4o e -8o , obtém-se

Resolvendo a equação de equilíbrio de momentos, obtém-se = -2,865485o

Wn

x

znWSqxxCCzCCzzCCcCz

T

FzwTDLFfxTxFwLDMAC tancossensencossencos

WFFn

FFF

x

zFMzFxFF

FnWF

SqCCFF

SqCCF

SqCCF

engaxx

Tzwzaz

T

FengACwwxwwzTz

axzeng

fxTxwxax

LDwx

DLwz

tancossen

sencos

sencos

Manobras Simétricas

altitude-4

velocidade-3

CG do posição e peso-2

carga defator -1

de depende

21

1 1

2

az

zzazazzaz

azazz

C

VSWn

SqWn

CqCWnF

W

F

W

Fn

Em vôo nivelado não-acelerado, o piloto puxa o manche

maior ângulo de ataque

maior fator de carga vertical

Hipótese: Velocidade constante.

Diagrama V-N de Manobras

20

2

0

22

02

2

2121

E

EazEazaz

zz

az

VV

W

SK

VKCW

VSC

W

VSCn

Sq

WnC

2Emáxazmáxz VKCn

Envoltória das condições permitidas para operação normal do avião.

velocidade verdadeira

nível do mar

velocidade equivalente

1sV

max1

01

2

nVV

SC

WV

sA

máxazs

Diagrama V-N de Manobras Flight maneuvering envelope (FAR Part 25§ 25.333)

Diagrama V-N de Manobras Design Airspeeds (FAR Part 25§ 25.335)

The selected design airspeeds are equivalent airspeeds (EAS).

VS0 = power off stalling speed at sea level, at the design landing weight, and in the landing configuration;

VS1 = stalling speed with flaps retracted;

VA = design maneuvering speed;

VB = design speed for maximum gust intensity;

VC = design cruising speed;

VD = design diving speed;

VF = design flap speeds;

VDD = design drag device speed.


(a). For VC, the following apply:

(1). The minimum value of VC must be sufficiently greater than VB 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), VC may not be less than VB + 1.32 UREF (with UREF as specified in § 25.341(a)(5)(i)). However VC need not exceed the maximum speed in level flight at maximum continuous power for the corresponding altitude.

(3). At altitudes where VD is limited by Mach number, VC may be limited to a selected Mach number.

§ 25.341(a)(5)(i). At the airplane design speed VC, positive and negative gusts with reference gust velocities (UREF) 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.


(b). VD must be selected so that VC/MC is not greater than 0.8 VD/MD, or so that the minimum speed margin between VC/MC and VD/MD is the greater of the following values:

(1). From an initial condition of stabilized flight at VC/MC, 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 § 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 MC 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.

(c). For VA, the following apply:

(1). VA may not be less than

where n is the limit positive maneuvering load factor at VC; and VS1 is the stalling speed with flaps retracted.

(2). VA and VS1 must be evaluated at the design weight and altitude under consideration.

(3). VA need not be more than VC or the speed at which the positive CNmax curve intersects the positive maneuver load factor line, whichever is less.

nVS1


AV

zn

HAAVgSV 1 EV

comNC max

incNC max

maxzn A

1 2A

(d).

(1). VB may not be less than where:

Kg = gust alleviation factor =

mg = airplane mass ratio =

VS1 = the 1-g stalling speed based on CNAmax with the flaps retracted at the particular weight under consideration;

VC = design cruise speed (knots equivalent airspeed);

Uref = 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.

= density of air (slug/ft3);

c = mean geometric chord of the wing (feet);

g = acceleration due to gravity (ft/sec2);

a = slope of the airplane normal force coefficient curve, CNA per radian;

(2). At altitudes where VC is limited by Mach number--

(i). VB may be chosen to provide an optimum margin between low and high speed buffet boundaries; and,

(ii). VB need not be greater than VC.


w

aVUKV Crefg

s 49811

gg 3.588.0

agcw 2

(e). For VF, 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). VF may not be less than-

(i). 1.6 VS1, with the flaps in takeoff position at maximum takeoff weight;

(ii). 1.8 VS1, with the flaps in approach position at maximum landing weight; and

(iii). 1.8 VS0 with the flaps in landing position at maximum landing weight.

(f). The selected design speed for each drag device (VDD) 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, VDD may not be less than VD. 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.


(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 VD 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 - 1.0 at speeds up to VC; and

(2). Must vary linearly with speed from the value at VC to zero at VD.

(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§ 25.337)

000,10

000,241.2

W

Fatores de Carga Típicos

Part 25 § 25.331 - 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 § 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). 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 § 25.397(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 VA. The airplane is assumed to be flying in steady level flight (point A1,§ 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 A2 in § 25.333(b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.

Part 25 § 25.331 - 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 § 25.337 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, § 25.333(b)). The positive acceleration must be equal to at least 39 n ( n – 1.5 ) / V Radians/sec2Where 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

Carga de Balanceamento da Empenagem Horizontal

yTTTxTTzFFengFFengACwwxwwz

zzFengazzFengTzwz

xxFengaxxFengTxwx

IMzFxFxFzFMzFxF

WngaWFFagWWFFF

WngaWFFagWWFFF

sencos

cossen cossen

sencossencos

wT

TwTTxFwFengFwFengACwz

T

TTTxFFengFFengACwwxwwz

BTz

xx

MzzFxxFzzFMWxn

x

MzFxFzFMzFxFF

sencos

sencos

cte

21

wTTw

EMCACA

wT

ACwzBTz

xxxx

VkcSqCM

xx

MWxnF

(FzT)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 MT e FxT(zT-zw), obtém-se

zEBTz nkVkF 22

1

Carga de Balanceamento da Empenagem Horizontal

wzEBTz xnkVkF 22

1

CA do atrásCG 0 ) wxb

nz

VE

A D

EH

CA do frente àCG 0 ) wxa

(FzT)B

H

E

D

VE

A

VE

A

D

E

H

(FzT)B

Distribuição de (FzT)B na Empenagem Horizontal

Tz

az

azzaz

TT

SqWn

C

SqCWnF

i

pT estabilizador profundor

Distribuição típica de pressão

p = 10o

T = 0o

p = 0o

T = 10o

Características aerodinâmicas da empenagem

TzTTzBTz

TTzTzTz

CqSCF

qSCSqCF

p =

0o

10o20o -10o

-20o

T

C´zT

p

Manobras não Corrigidas(Unchecked Maneuvers)

nz

VE

AD

EH

nz =1 A1

A2

VA

(1). Maximum pitch control displacement at VA. The airplane is assumed to be flying in steady level flight (point A1,§ 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 A2 in § 25.333(b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.

zn

TL

t

p

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 § 25.337 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, § 25.333(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

nz

VE

A D

EH

nz =1 A1

A2

VA

D2

D1

t

znp

TL

Manobras não Corrigidas – Análise Aproximada

x

z

BTx

BTzF

BTzBTz

zazAz

CF

SqWn

CVVn

e 1 T

nz

VE

AD

EH

nz =1 A1

A2

VA

yy

BTBTzMTMTzy

TMTzMTz

IM

xFxFM

qSCF

MTx

MTzF

x

z

Antes da manobra

Depois de p

p =

0o

10o20o -10o

-20o

T

C´zT

BTzC A1

MTzC A2

Manobras Corrigidas – Análise Aproximada

x

z

Tx

BTzF

T

yTz x

IF

nz

VE

A D

EH

nz =1 A1

A2

VA

D2

D1

x

z

Tx

BTzF

T

yTz x

IF

Carga para baixo

[nz = 1; = 39 nmax (nmax - 1) / V ; V dado em knots]

Carga para cima

[nz = nmax; = -26 nmax (nmax - 1) / V - V dado em knots]

Rajadas

Rajadas de contorno vivo

Rajadas com gradiente

FAR Part 23 §23.341

Diagrama V-n de Manobras e Rajadas

FAR Part 25 §25.341

Efeitos de rajadas na empenagem horizontal

Rajadas de Contorno Vivo

Rajadas de Contorno Vivo

EEEE

EEEREE

VUVU

VUVVUV

arctg

22

12

EEzz VUSW

ann 02

1

11

Variação instantânea:

EEeE

Ewzwz VUSaVS

VU

aqSddCF 02

0 21

21

EEwz

z VUSW

a

W

Fn 02

1

VE

UE

VR

a- o deslocamento vertical do avião reduz o efeito da rajada e a rotação em arfagem;

b- efeito de Wagner: retardo na variação de Czw 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 Kg

Rajadas com Gradiente (Graded Gust)

SW

VUaSKn

VUaSKF

EEgz

EEgwz

0

0

21

21

EgE UKU

Rajadas (FAR Part 23 §23.341)

SW

aVUKn Edeg

498

g

ggK

3,5

88,0 gac

SW

og

2

Ude = Derived gust velocities referred to in § 23.333(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 CNA 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 CL 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.

§ 23.333(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 VC 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 VD 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 VB 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.

Diagrama V-n de Manobras e Rajadas

FAR Part 23 §23.333

Válido para um dado peso (W) e posição do CG do avião

(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 musttake 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)


H

sUU de

cos12

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 = UrefFg(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.

(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)

Rajadas (FAR Part 25 §25.341)(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.

weight takeoffmaximum

weightfuel zero maximum

weight takeoffmaximum

weightlanding maximum

4tan 000,2501

2

1

12

R

R

RRFZF gmmogz

(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.

Efeitos de Rajadas na Empenagem Horizontal

aC

d

Cd

T

Tz

T

Tz

qSaqSCF TTTTzTz

SVaUK

F

V

UK

EEg

wz

E

Eg

02

C´zT

T

T

C´zT

TEEg

ETE

EgTz

E

EgTT

TTT

SVUdd

aK

VSVU

dd

aKF

V

UK

dd

dd

dd

dd

i

0

20

12

21

1

1

1

Na asa: Na empenagem:

No avião:

yTTxwwxTTzwwzy

TEE

gTzwzaz

IzFzFxFxFM

SS

dd

aaSVUK

FFF

1

2 0

FAR Part 25 §25.343 – 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 § 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 conditions of § 25.341(a) but assuming 85% of the design velocities prescribed in § 25.341(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: § 25.1001(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.

FAR Part 25 §25.345 – 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 § 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-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 § 25.341(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.

Efeitos Estruturais das Manobras Simétricas

Ponto A: alto ângulo de ataque positivo

nz

VE

A D

EH

L

D

Fzw

Fxw

Ponto D: baixo ângulo de ataque positivo

L

D

Fzw

Fxw

(-) (-) (-) (-) (-) (-)(+) (+) (+) (-) (-) (-)

(+) (+) (+) (-) (-) (-)(+) (+) (+) (+) (+) (+)

MFzw

MFxw

críticos

MFzw

(-) (-) (-) (-) (-) (-)(-) (-) (-) (+) (+) (+)

(-) (-) (-) (+) (+) (+)(+) (+) (+) (+) (+) (+)

MFxw

críticos

Efeitos Estruturais das Manobras Simétricasnz

VE

A D

EH

Torção

Ao longo de AD variam Fzw e Fxw, embora Fza = nzW = cte. As forças de inércia, que dependem de nz, são constantes, de modo que a flexão na asa varia ao longo de AD.

yy zEy

zwzzTzwzaz

ECA

zwzCAy

nkVkM

nkFWnFFF

VkM

WbnaFMM

22

1

21

My máximo positivo para VE mínimo e nz máximo positivo

→ Ponto A do diagrama V-n

My máximo negativo para VE máximo e nz máximo negativo

→ Ponto E do diagrama V-n

CACG

CT

Fzw

Fxw

MCA

a b

nxW

nzW

My

Flexão nas asas

Variação das Cargas com o Peso e Posição do CG

wT

wzCATz xx

WxnMF

CGCA

CAT

nzW

Fzw

MCA

FzT

xw

xT

máximo) o que domenor peso com ente(eventualm

CG do frente à mais posição e , para dado é máxzmáx nWM máx

00

0

Tzw

CA Fx

M

nzWf

FzT

Ex. Esforços Máximos em Seção da Fuselagem Traseira

Rajada com Peso Mínimo

SW

VUaKn

VSUaK

F

FFF

EEgz

EEg

wz

wzwzwz

21

2

0

0

rajada da antes

Fzw é máxima para CG à frente e Wmax

nz é mínimo para Wmax

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)

Carga Variável na Asa

iz

izz

A

MWWn

MWnWn

M

11

2

21

lll

ll

comb

comb

2

22

AA MM 12 ll se

Flexão pode ser crítica com peso abaixo do máximo

a) Com combustível

Mi = contribuição dos demais itens de inércia da asa

b) Sem combustível

combustível

l1 l2nzWcomb/2

nzW/2A

A

Peso do avião = W - Wcomb iz

A MWWn

M 1lcomb2

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

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