Turbine film cooling technology is widely used in high thrust-weight ratio aero engine nowadays, so the accuracy in manufacturing of turbine film cooling holes will directly affect the performance of the engine. The film cooling micro-hole has the characteristics of small diameter, high quantity, high depth-diameter ratio, and complex spatial angle, which has a complicated fabrication processes and high precision requirements. The drilling technology and equipment for film cooling holes abroad are strictly kept secret, while, there is still a fact that the low forming and unsteady quality of film cooling holes produced domestically. Therefore, it is of necessity to make a summary of the current situation and development trend of producing film cooling holes for turbine blade to provide a reference for the development of film cooling hole manufacturing technology and its equipment. In this paper, the necessity and importance of the film cooling hole accuracy are summarized and the state-of-the-art manufacturing technologies for producing drilled cooling holes are analyzed, and then, the research content and the solution to the key problems for the film cooling hole drilling technology are proposed based on the analysis. Moreover, in view of the need for precise measurement of the film cooling holes, the existing film cooling hole measurement technol  y is introduced. Finally, the development trend of drilling and measuring film cooling hole of the hollow turbine blade is generalized according to their technical status and key problems.

A350 aircraft and Boeing787 aircraft are the most advanced representative of modern civil aircraft,which have the relative higher composite material content. The special design forms a large number of hybrid structures between the metal and composite materials such as wing body butt structure. It is of practical significance
to study the docking structure of the two models. The design principle on selection range of the wing body docking process interfaces of the civil aircraft is described. The differences and characteristics of the two types of docking process interface are illustrated, and the structures of the two types of aircraft wing panel butt, the wing beam butt and the outer wing panel-fuselage panel butt are analyzed, and the advantages and disadvantages of connection structures of the two aircrafts above are analyzed and compared in this paper. The influence of the two docking structures on wing assembly and final assembly is explained, and the domestic research status and corresponding design ideas of the hybrid structure connection are expounded. Finally, the characteristics of the docking structure of the two models are summarized.

In order to solve the problem of low accuracy of the fuselage of large aircraft in China, this paper studies　the docking method on the fuselage of a large aircraft. First of all, this paper selects and designs the docking datum of the fuselage and the docking station of the fuselage. Later, in the process of the fuselage docking, using a laser collimator, make the laser collimator light polymerization through artificial adjustment, and the desired data displayed on the digital display table, so as to adjust the body of the aeroplane. Again measuring actual coordinates of fuselage assembly datum using laser tracker, online measurement control system trims fuselage according to the actual coordinates measurement data to meet the requirements of the position of the fuselage and docking. Finally, according to the fine-tuning of the fuselage, the actual sitting value of the measured reference point and the square root of the sum of distance squares of the theoretical coordinates
are used to determine whether the fuselage butt meets the standard. The method of fuselage docking in this paper can reasonably control the deviation at the tolerance range. Through the study of large aircraft fuselage docking technology, laser tracker, laser collimator, online measurement software and servo drive device as the main body of online measurement control system, can greatly improve the accuracy of large aircraft fuselage docking.

In order to solve the problem of the positioning and nondestructive holding of the fuselage panel of the large composite material, the paper puts forward the flexible tooling structure scheme based on mixed joint positioning and vacuum adsorption. On the basis of the analysis of the position process and the vacuum adsorption process of composite materials panel, design a kind of integrated X axis and Y axis tandem cross sliding table, parallel connection of Z axis and B angle adjustment module, vacuum adsorption and A angle adaptive flexible tooling structure. And the function test results of flexible tooling test pieces show that the positioning and adsorption of the flexible tooling are stable and accurate, and the adsorption is safe and reliable, which satisfy the requirements of positioning and nondestructive clamping function.

It is widely acknowledged that some special drilling areas exist in aircraft assembly. For example, the wing root and the vertical tail. Not only are they narrow in space but also require certain height, which makes it difficult to fully harness the advantages, especially in the usual automatic drilling device and robot drilling system. Hence, a portable helical trajectory drilling device was designed in order to solve the problems. This device is compact, portable and could set process parameters according to the program, besides, it may be applied in drilling of aeronautical materials. Moreover, tool feed speed can be changed automatically for multilayer material. In this paper, the authors applied the developed spiral trajectory drilling device to examine the drilling of 7075–T7351 multilayer Al alloy material. It is showed that the accuracy of this device reached the expectation and the rationality of the design and control methods of the new type drilling device was verified.

Spin riveting is a kind of riveting methods, in which the rivet shaft presses the local of rivet, and swings continuously around a central pivot until rivet shaped. After forming, the rivet material is continuous without breaking, bending, bulging, and the parts connected with the rivet are not deformed. In the 2017 materials as the object of ϕ6 rivets, based on Simufact. Forming software, using the method of tracking its spin riveting and the numerical simulation on forming process, extracting the stress values of rivets during certain points of the deformation zone, and the data is transformed to the equivalent stress-travel curve, then explore deformation mechanism of spin riveting. According to the structure of the wind panel and its assembly process, the deformation mechanism of the spiral riveting is applied to the rotary riveting equipment. By optimizing the structure of the riveting equipment, the CNC technology is introduced into the equipment to realize automatic processing and improve the efficiency and quality of wall panel processing.

To meet requirements of measure during aircraft assembly process, such as big dimension, high accuracy, multitask and quick measure, the methods are developed including precision calibration of measure basis network facing on aircraft digital assembly, multi-systems cooperative measure, and technology solutions are described. Then measuring basis network covering total aircraft based on USMN is built to improve accuracy measuring reference. Work pattern of cooperative measure including multi-system is established to extend measuring scope and accuracy. The method was developed in this paper can realize the measurement of internal hidden features, hole position information, axis, mating surface, aerodynamic profile, flush and gap on the aircraft skins, meets the requirements of three-dimensional digital measurement for aircraft assembly. At last, the methods of cooperative measure are validated by measure process experiment of huge components assembly and huge parts surface feature including aerodynamic profile, flush and gap on the aircraft skins.

As the key link of aircraft development, aircraft assembly has many characteristics, multi-disciplinary field, high accuracy of coordination, high labor intensity and so on. The advanced aircraft assembly technology can improve efficiency of aircraft assembly, shorten development cycle, and lower the production cost. Now most domestic aviation companies still use non-modular assembly. They don’t form a high efficient, neat and more utilizable assembly schema. To analyze non-modular aircraft assembly way, advanced modular aircraft assembly concept is proposed, The focus is to describe several key technology, modular design, the definition of assembly module, the construction of the assembly structure tree, the management of the assembly order library, based on product configuration matching assembly order.

In view of the engineering realization of the modular assembly of a large aircraft wing, the technical background and the modular features of the aircraft components module are analyzed, and the module assembly process and the influencing factors are combed. The engineering design of wing assembly process, wing integration / flexible assembly and active airfoil unit is carried out. The integrated flexible working platform and rail mounting vehicle are designed. The design and analysis of control logic structure, test method and test equipment planning for aircraft active airfoil are finished. Finally, the application of wing module assembly and its effect are introduced.

The distribution of residual stress field in perforated disks is studied through numerical simulation method. Because of the limitation of current measurement technology, the feasibility of using 3D finite element model for numerical simulation is analyzed. The results show that the degree of cold expansion (DCE) has the greatest effect on the area of the wall near the upper surface. Furthermore, the overlarge DCE will make the tangential compressive residual stress field in this area no longer exist. It is also found that TC4 titanium alloy single-hole circular plate is more suitable for the cold extrusion process than 7050–T7451 aluminum alloy perforated circular plate.

To manufacture the aircraft frame and rib parts rapidly and precisely, the approach of intelligent manufacturing technology development is proposed from the viewpoints of knowledge base and manufacturing model, and the intelligent manufacturing flow is established. The scheme from technology analysis, implementation to application is designed according to the application target. The technology has been applied in AVIC corporations of Hongdu, Shenyang and Chengdu. The test result shows that the forming accuracy of outer shape, flange angle and height is improved significantly and precise forming is achieved. The technology can be further applied in the future.

Stretch forming is the basic method of aircraft sheet metal forming. Because the ultra-thick skin is easy to be tiled during the stretching process, the finite element model is established based on the analysis of the trajectory design and the resilience stress in this paper. Through the analysis of the super-thick double convex skin, the pretensioning–covering–pull-up loading trajectory is the most suitable, and can effectively guarantee the fitting degree of the parts. At the same time, combining with the finite element analysis of displacement control, pressure control and springback analysis set clamping spacing at 22t, clamping pressure at 7mm, the distribution of stress, amount of thinning and the springback amount of uniform distribution, which experiments that the parameters can effectively control tire degree, and achieve the design requirements.

Technical problems existing in molding process of U-shaped sandwich structure composite skin was studied. Many aspects of molding process were ameliorated, which included honeycomb core stabilization process, honeycomb core shape control, cross interval grasped layer and shape detection of parts. The results showed that stabilization process for honeycomb core could make greatly improve the quality of honeycomb core chamfer of manual cut. The optimal design of shape models could great assistance in shape control of honeycomb core. The cross interval grasped layer could avoid pressure conduction problems existing in molding process, and effectively solve the problems of honeycomb shrinkage, layer folds and internal delamination. The association, which included simple snap-gauge, U-shaped clamping devices and feelers, realized the efficient measurement of the shape of the parts.

In this paper, U-shaped components of carbon fiber composite with two different-height flange slabs are studied. Autoclave moulding process and VARIM process are used to prepare the U-shaped components. The U-shaped component angular contraction conditions are studied by measuring their angle. And based on the measurement results, the mould compensation angle is optimized and redesigned. The angular contraction conditions of the U-shaped components made by the two above-mentioned processes are different. When designing the compensation angle of the U-shaped component mould, it is impossible to generalize all the moulds used in different molding processes and we should treat them differently according to the molding process characteristics. In addition, for the U-shaped components with two differentheight flange slabs, whichever kind of molding process they are made by, the degree of angular contraction on the high flange slab side is smaller than that of on the lower flange slab side. When designing the compensation angle of the U-shaped component mould, this distinction should not be neglected.