A coupling device for rotably coupling a shaft with a gearbox in a geared turbo fan aircraft engine, wherein the coupling device includes a connection to the shaft at a first end and a connection to the gearbox at a second end, the first and the second ends being axially separated and at least one curved shape between the first end and the second end extending from the gearbox radially inwards to the shaft and the at least one curved shape including at least one cross-section in the axial direction of the engine with a logarithmic profile or a power profile.

An apparatus comprises a gas turbine engine having a compressor for increasing a total pressure of a working fluid prior to delivery to a combustor for mixing with fuel. The compressor has a plurality of airfoil members that extend between a first flow surface and a second flow surface. The plurality of airfoil members have a first camber shape at a tip portion that provides a cumulative integral according to the relationship $0.4<\frac{{\int }_0^1\left[{\int }_0^x{\beta }^{\prime }\mathrm{dx}\right]\mathrm{dx}}{{\int }_0^1{\beta }^{\prime }\mathrm{dx}}\le 0.5$
where x is chord location, ${\beta }^{\prime }=\frac{{\beta }_i-\beta }{{\beta }_i-{\beta }_e},$
β.sub.i is inlet metal angle, β.sub.e is exit metal angle, and β is a camber angle at a given chord location.

A pneumatic device includes an outer ring (1) and a core body (3), at least one stage of secondary stroke flow channel (300) being provided between a nozzle (301) and an exhaust port (302) which are located at an outer ring surface of the core body (3); gas enters from an intake passage (31), is ejected in stages through the nozzle (301) and the secondary stroke flow channel (300) of the core body (3), acts on at least two driving recesses (11) in a circumferential direction of the outer ring (1), and generates a pushing force for the driving recesses (11) to push the outer ring (1) to rotate and do work, so as to achieve a power output, and finally, the gas is discharged from an exhaust passage (310) through the exhaust port (302) of the core body (3).

A pneumatic device includes an outer ring (1) and a core body (3), at least one stage of secondary stroke flow channel (300) being provided between a nozzle (301) and an exhaust port (302) which are located at an outer ring surface of the core body (3); gas enters from an intake passage (31), is ejected in stages through the nozzle (301) and the secondary stroke flow channel (300) of the core body (3), acts on at least two driving recesses (11) in a circumferential direction of the outer ring (1), and generates a pushing force for the driving recesses (11) to push the outer ring (1) to rotate and do work, so as to achieve a power output, and finally, the gas is discharged from an exhaust passage (310) through the exhaust port (302) of the core body (3).

A coupling device for rotably coupling a shaft with a gearbox in a geared turbo fan aircraft engine, wherein the coupling device includes a connection to the shaft at a first end and a connection to the gearbox at a second end, the first and the second ends being axially separated and at least one curved shape between the first end and the second end extending from the gearbox radially inwards to the shaft and the at least one curved shape including at least one cross-section in the axial direction of the engine with a logarithmic profile or a power profile.

An apparatus comprises a gas turbine engine having a compressor for increasing a total pressure of a working fluid prior to delivery to a combustor for mixing with fuel. The compressor has a plurality of airfoil members that extend between a first flow surface and a second flow surface. The plurality of airfoil members have a first camber shape at a tip portion that provides a cumulative integral according to the relationship

[00001]$0.4<\frac{\underset{0}{\overset{1}{}}.Math.\left[\underset{0}{\overset{x}{}}.Math.{}^{}.Math.\mathrm{dx}\right].Math.\mathrm{dx}}{\underset{0}{\overset{1}{}}.Math.{}^{}.Math.\mathrm{dx}}0.5$

where x is chord location,

[00002]${}^{}=\frac{{}_i-}{{}_i-{}_e},$

.sub.i is inlet metal angle, .sub.e is exit metal angle, and is a camber angle at a given chord location.

An apparatus comprises a gas turbine engine having a compressor. The compressor airfoil member is disposed in the compressor and has a camber angle variation along a chord and a concave pressure side surface at a tip portion of the compressor airfoil member from a leading edge to an intermediate-chord location. This first camber shape results in a normalized camber angle that reaches at least 0.4 in the first 10% of normalized chord, and the normalized camber angle is defined as ${}^{}=\frac{{}_i-}{{}_i-{}_e}$
where .sub.i is inlet metal angle, .sub.e is exit metal angle, and is a camber angle at a given chord location.

Rockets, rocket motors, methods of controlling a rocket and methods of evaluating a rocket design are disclosed. In some embodiments, a method of controlling a rocket may include measuring a combustion chamber pressure, calculating a logarithm of the measured combustion chamber pressure, and computing the difference between the logarithm of the measured combustion chamber pressure and the logarithm of a reference combustion chamber pressure value to generate an error signal. The method may further include filtering the error signal to generate a compensated signal in the logarithm domain, and exponentiating of the compensated signal in the logarithm domain to provide a compensated signal in the physical domain.