In this study, the aerodynamic performances and the flow separation behaviour of the wing profiles, used in the wing of Boeing 737-Classic aircraft, and a model wing formed of these profiles were investigated at 2 × 105 Reynolds number (Re) and different angles of attack ranging from −4° to the angle, which maximum lift point obtained. The experiments were conducted in a low-speed wind tunnel in Erciyes University Faculty of Aeronautics and Astronautics Aerodynamic Laboratory. Four profiles and the model wing, which was obtained from the combination of the four profiles were produced in a rectangular shape with dimensions of 0.21m × 0.21m. In the first part of this two-part study, the wing profiles were examined individually regarding aerodynamic performance and flow separation. In the second phase of the study, the aerodynamic performance and the flow separation behaviour of the model wing were examined and the results were compared with the data obtained from each profile. The study results showed that the slopes of the numerical graphics are generally quite consistent with experimental results. In addition, the pressure and velocity distributions have followed the normal trend until the angle-of-attack (AoA) of 20°. In contrast, the negative speed values and negative pressure zones have appeared on the profile above this angle. The air flowed as laminar on the profiles and the model wing until 20°, while laminar separation bubbles are begun to take place at about 24° AoA. Finally, the best aerodynamic performance has been obtained with the model wing.

As concern for the environment increases, so does the desire to reduce emissions and improve fuel efficiency. One avenue being investigated in the aerospace industry is the reduction in losses in a gas turbine engine through endwall contouring, which has shown some promise in improving efficiency. The current investigation was aimed at investigating the unsteady effects downstream of a rotor with a non-axisymmetric endwall contour originally designed for a cascade.

Previous investigations (numerical and experimental) have shown that the endwall contour improved the turbine efficiency by reducing the cross passage secondary flows. The contour however had a detrimental effect on the tip flow. The current investigation found that there were negligible unsteady differences between the annular and contoured rotor, with the average comparing well to the steady state results. It was also found that the effects at the tip were due to the experimental blades untwisting during operation.

We previously described an efficient, lightweight and flexible electro-thermal system, based on directly drawn carbon nanotube web (CNT web), as part of an icing protection system for carbon fibre reinforced polymer (CFRP) composite aircraft structures. The location of the heating elements on critical lifting surface leading edges or nacelle intake lips makes them particularly susceptible to impact damage, which may leave no visible mark. This makes it desirable to have both a mechanism for identifying the location of damage to the CNT structure (and by inference, potential damage to the underlying CFRP) and a process for restoring the CNT heater to full operation. With the CNT web acting as a sensor, impact damage is identified by an increase in electrical resistance and, particularly, by infrared imaging, which reveals a cold spot or zone depending upon the CNT web layup. Whereas a unidirectional CNT web layup exhibits a large increase in resistance and loss of a full width band of operation, a cross ply quasi-isotropic CNT web arrangement suffers only a small increase in resistance and a loss of function that is highly localised to the damaged area. A novel methodology, based on dispersed CNT in resin, is described for repairing and reconnecting the CNT structure and restoring functionality. A CNT web-based electro-thermal element was applied to the leading edge of a representative carbon-fibre composite wing section to demonstrate the flexibility of this system.

Variable camber flap technology can adjust the spanwise circulation distribution, thereby reducing the induced drag. Therefore, the concept of variable camber flap is introduced into the design of propeller aircraft wing, and the design for drag reduction of propeller aircraft is carried out. The numerical simulation of the propeller aircraft is carried out by using the actuator disc method with non-uniform distribution of radial and circumferential loads. Through the unsteady simulation of a single propeller, the aerodynamic load on a periodic propeller is extracted as a boundary condition to the steady simulation of the full aircraft. The load extracted by the actuator disc is compared with the unsteady simulation result, which verifies the reliability of the method. The design for drag reduction at cruise and climb design conditions are respectively carried out with the variable camber flap technology. The variable camber cruise configuration is evaluated at both the begin and end cruise conditions. The results show that, after the flaps deflecting at a small angle according to the circulation distribution, the camber distribution of the wing is adjusted to make the circulation distribution closer to the elliptical circulation distribution. At the design cruise condition, the drag coefficient is reduced by 1.4 counts, and the lift-drag ratio increase by 0.1. At both begin and end cruise conditions, the drag coefficient decreases by 1 count, and the lift-drag ratio increases by 0.07. At the design climb condition, the drag coefficient decreases by 1 count, and the lift-to-drag ratio increases by 0.09.

In a recent series of papers, Poll and Schumann have been developing a simple model for estimating fuel burn for turbofan powered, civil transport aircraft for a given mass, Mach number and flight level and in a specified ambient temperature profile for all phases of flight. This paper focuses upon the combination of Mach number and flight level at which an aircraft cruises with the absolute minimum fuel burn. For a given aircraft type, the information necessary to determine these conditions must be specified and this poses a challenge. An initial attempt to obtain these data has been described previously by the first author. In this paper, the optimum conditions are found using a completely different approach. Starting from first principles and using established theory, the equations governing the situation where engine overall efficiency and airframe lift-to-drag ratio both have local maxima at the same flight condition are developed. This special case is termed the “design optimum” condition and, for a specified aircraft mass and a specified atmospheric temperature versus pressure profile, it gives the lowest possible fuel burn for any aircraft and engine combination. The design optimum occurs at a particular Mach number and Reynolds number, and it is a fixed characteristic of the aircraft. The analysis reveals the significance of Reynolds number variations, wave drag, including its derivatives with respect to both lift coefficient and Mach number, and the atmospheric properties. Whilst wave drag is notoriously difficult to determine accurately, it is found that solutions to the equations are not particularly sensitive to the accuracy of this quantity. Consequently, a simple, physically realistic model can give good results. An appropriate model is developed and a complete, approximate solution is obtained. Taking the International Standard Atmosphere as the design atmosphere, results are presented for the 53 aircraft types previously considered by Poll and Schumann. Relative to the design optimum conditions, when Reynolds number is constant and wave drag is zero, compressibility alone reduces L/D by about 5%, reduces lift coefficient by about 1.5% and increases drag coefficient by about 3.5%. Reynolds number variation has little effect upon L/D, but it reduces lift coefficient and drag coefficient by a further 7% and 8% respectively. The reduction in lift coefficient has a significant impact on the optimum cruise flight level.

In general, an aircraft’s operating optimum will not coincide with its design optimum, but deviations are expected to be small. Therefore, using the design optimum solution as a reference point, an improved version of the operating optimum estimation method described by Poll and Schumann in previous work is developed. This allows the estimation of the conditions for absolute minimum fuel burn for an aircraft of given mass flying thorough any atmosphere. Updated coefficients for the 53 aircraft types are given.

Coupling of clearance joint and harsh aerodynamic heating environment is an inevitable nonlinear factor in folding mechanism of the fin of high-speed aircrafts that remarkably modifies natural frequencies and modes of vibration from the initial design state. However, accurately predicting dynamic properties of deployable fin with full consideration of these effects is not common industry practice. A practical semi-analytical model based on Hertz contact theory and ESDU-78035 model is proposed in this study to investigate high-temperature connection stiffness of local hinged–locked mechanisms. Material property degradation and clearance variation caused by thermal expansion are comprehensively considered and quantified in this model. Vibration characteristics of the assembled deployable fin are then solved using finite element method (FEM). The real-time evolutionary process of thermal mode of the fin is discussed. And natural frequencies of fixed-value and time-varying connection stiffness are compared. The simulation results of this study demonstrate that the relative error of structure temperature between the sequential approach and fully coupled simulations is less than 6.98%. The connection stiffness (slope of the load-displacement curve) of the folding mechanism under high temperature conditions decreases by 3.52%, and the variation is mainly caused by the degradation of the elastic modulus of the material, while the clearance change due to the thermal expansion has no significant effect on the slope. The natural frequency of the deployable fin exhibits an inverse correlation with the temperature change trend, and the first three frequencies decrease by 1.67, 7.75, and 16.28 Hz compared to the initial value, respectively.

The overall efficiency of a turbofan engine may be expressed as a function of the Mach number, flight level and one other parameter. This may be either the net thrust, the turbine entry temperature or the fuel flow rate. Using basic aero-thermodynamic principles, dimensional analysis, normalisation and curve fitting, five approximate and “near universal” relations have been identified for engines having bypass ratios between 1 and 13. These relations contain five independent characteristic engine parameters. When these parameters are known, the relations form the basis of an estimation method for engine overall efficiency that is simple, fast, open source, completely transparent and, as new information appears, capable of further refinement. Since the empirical relations presented in this analysis are valid for Mach numbers greater than 0.2, the method is applicable to all airborne phases of flight. For a given aircraft, if the flight trajectory is specified in sufficient detail for the variation of net thrust with Mach number and flight level to be determined, only three of the five relations, together with the value of engine overall efficiency at a single reference condition, are needed to estimate the overall efficiency at every point on the trajectory. Comparisons with the data used in this analysis suggest that the accuracy is better than ±5% in most cases. In the completely general case, two additional engine characteristic parameters, one a total temperature ratio and the other a Mach number, are introduced. If these are known, both engine overall efficiency and net thrust can be expressed as functions of Mach number, flight level and turbine entry temperature. This allows the method to be used for the estimation of operating limits in the various phases of flight and in simplified optimisation studies, e.g. finding the environmentally optimum flight trajectory.

In previous work, estimates of engine overall efficiency at the “design optimum” condition have been estimated for 53 aircraft and engine combinations. It is shown that the ‘design optimum’ condition is an appropriate choice for the engine reference condition. Updated and revised values for the relevant parameters for these 53 examples, together with estimates for the two additional engine characteristic parameters, are given in tabular form.