Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, June 1997. / Thesis advisor, Raymond P. Shreeve. AD-A333 428. Includes bibliographical references (p. 49-50). Also available online.

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The existing 3-stage research turbine at Turbomachinery Performance and Flow Research Laboratory (TPFL) facility, Texas A & M University (TAMU) was replaced with a newly designed and manufactured 2-stage turbine in accordance with the design requirements as per DooSan, DHI. This new design of turbine consisted of bowed stator and rotor blades to study the effect on reduction of secondary ow losses and thus improvement in turbine efficiency if any. The new design also incorporated labyrinth seals on both inner and outer shrouds. Extensive Inter-stage and Performance experiments were carried out on this new turbine. Inter-stage measurements were accomplished by traversing three 5-hole probes radially and circumferentially, using the existing probe traverse system in TPFL. Performance tests were conducted for varying pressure ratio, at fixed rotational speed and for varying rotational speed with fixed pressure ratio and the efficiency was plotted against u/c_0. Each condition was tested and measured two to three times to check for reproducibility of the data. The results from inter-stage experiments show that the rotor row loss coefficient is about four times higher than the stator row loss coefficient. This high rotor loss coefficient reduces the total to static efficiency. From the performance tests, the maximum total-to-static efficiency observed was 85.2 percent located at around u/c_0 = 0.75. This relatively low efficiency is in consonance with the inter-stage results (high rotor loss coefficient).

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Helicopter rotor system operates in a highly dynamic and unsteady aerodynamic environment leading to severe vibratory loads on the rotor system. Repeated exposure to these severe loading conditions can induce damage in the composite rotor blade which may lead to a catastrophic failure. Therefore, an interest in the structural health monitoring (SHM) of the composite rotor blades has grown markedly in recent years. Two important issues are addressed in this thesis; (1) structural modeling and aeroelastic analysis of the damaged rotor blade and (2) development of a model based rotor health monitoring system. The effect of matrix cracking, the first failure mode in composites, is studied in detail for a circular section beam, box-beam and two-cell airfoil section beam. Later, the effects of further progressive damages such as debonding/delamination and fiber breakage are considered for a two-cell airfoil section beam representing a stiff-inplane helicopter rotor blade. It is found that the stiffness decreases rapidly in the initial phase of matrix cracking but becomes almost constant later as matrix crack saturation is reached. Due to matrix cracking, the bending and torsion stiffness losses at the point of matrix crack saturation are about 6-12 percent and about 25-30 percent, respectively. Due to debonding/delamination, the bending and torsion stiffness losses are about 6-8 percent and about 40-45 percent after matrix crack saturation, respectively. The stiffness loss due to fiber breakage is very rapid and leads to the final failure of the blade. An aeroelastic analysis is performed for the damaged composite rotor in forward flight and the numerically simulated results are used to develop an online health monitoring system. For fault detection, the variations in rotating frequencies, tip bending and torsion response, blade root loads and strains along the blade due to damage are investigated. It is found that peak-to-peak values of blade response and loads provide a good global damage indicator and result in considerable data reduction. Also, the shear strain is a useful indicator to predict local damage. The structural health monitoring system is developed using the physics based models to detect and locate damage from simulated noisy rotor system data. A genetic fuzzy system (GFS) developed for solving the inverse problem of detecting damage from noise contaminated measurements by hybridizing the best features of fuzzy logic and genetic algorithms. Using the changes in structural measurements between the damaged and undamaged blade, a fuzzy system is generated and the rule-base and membership functions optimized by genetic algorithm. The GFS is demonstrated using frequency and mode shape based measurements for various beam type structures such as uniform cantilever beam, tapered beam and non-rotating helicopter blade. The GFS is further demonstrated for predicting the internal state of the composite structures using an example of a composite hollow circular beam with matrix cracking damage mode. Finally, the GFS is applied for online SHM of a rotor in forward flight. It is found that the GFS shows excellent robustness with noisy data, missing measurements and degrades gradually in the presence of faulty sensors/measurements. Furthermore, the GFS can be developed in an automated manner resulting in an optimal solution to the inverse problem of SHM. Finally, the stiffness degradation of the composite rotor blade is correlated to the life consumption of the rotor blade and issues related to damage prognosis are addressed.

Thesis (MScEng)--Stellenbosch University, 2011. / ENGLISH ABSTRACT: Turbomachines are exposed to several environmental factors which may cause failure of components. One of these factors, high cycle fatigue, is often caused by blade utter. This thesis forms part of a project of the European Seventh Framework Programme (FP7), called project Future. Project Future is doing theoretical and experimental investigation into the occurrence of utter in turbomachinery. The objective of this thesis was to evaluate the effectiveness of a gas injection system as a means of exciting vibrations on the rst stage rotor blades of a compressor. Unsteady simulations of the excitation velocity perturbations were performed in the Computational Fluid Dynamics (CFD) software, Numeca FINE/Turbo. Experimental testing on the in-house Rofanco compressor test bench, using one prototype of the 15 injector system, provided data that was used to implement boundary conditions and to verify certain aspects of the unsteady simulation results. The simulation results revealed the following: the injector bypass frequency was so dominant that the excitation frequency was hardly detectable in the majority of cases. Furthermore, several secondary frequencies were consistently present. The injector bypass frequency, as well as the secondary frequencies, occurred as a result of the convolution of Fast Fourier Transforms. While the injector bypass frequencies can theoretically be eliminated, it will not be possible to eliminate the secondary frequencies from the blade response. In conclusion, according to the CFD results, it will not be possible to excite a single excitation frequency by making use of a nite number of gas injector vibration exciters. / AFRIKAANSE OPSOMMING: Turbomasjiene word onderwerp aan verskeie omgewingsfaktore wat falings van komponente kan veroorsaak. Een van hierdie faktore, naamlik hoëfrekwensie vermoeidheid, word onder andere veroorsaak deur lem adder. Hierdie tesis is deel van 'n projek in die Sewende Europese Raamwerk Program (European Seventh Framework Programme - FP7), projek Future. Projek Future doen teoretiese en eksperimentele ondersoek na die voorkoms van lemfl adder in turbomasjienerie. Die doelwit van hierdie tesis was om die effektiwiteit van 'n gasinspuiter vibrasie-opwekkingstelsel te evalueer, deur gebruik te maak van onbestendige simulasie in die berekenings vloei-meganika sagtewarepakket, Numeca FINE/Turbo. Eksperimentele toetswerk op die plaaslike Rofanco kompressortoetsbank, met 'n prototipe van die 15 inspuiter stelsel, het inligting verskaf wat gebruik is om die inlaattoestande te spesi seer en simulasieresultate te korreleer. Die simulasieresultate het getoon dat die frekwensie waarteen 'n lem by die inspuiters verbybeweeg, so prominent is, dat dit in die meerderheid van gevalle baie meer prominent is as die opwekkingsfrekwensie. Verder was daar ook deurgaans 'n aantal sekondêre frekwensies teenwoordig. Die teenwoordigheid van die inspuiter verbybeweeg frekwensie en die sekondêre frekwensies is die resultaat van die konvolusie van Vinnige Fourier Transforme. Alhoewel dit in teorie moontlik sal wees om die inspuiter verbybeweeg frekwensie te elimineer, is dit onmoontlik om die sekondêre frekwensies uit die lem vibrasie te elimineer. Ter opsomming, volgens die berekenings vloei-meganika resultate, is dit nie moontlik om met 'n stelsel van 'n eindige aantal inspuiters, 'n enkele vibrasie frekwensie op te wek nie.

For optimal rotor performance in a tiltrotor aircraft the difference in the inflow and the rotor speeds between the hover and cruise flight modes suggests different blade twist and chord distributions. The blade twist rates in current tiltrotor applications are defined based upon a compromise between the figure of merit in hover and propeller efficiency in airplane mode. However, when each operation mode is considered separately the optimum blade distributions are found to be considerably different. Passive blade twist control, which uses the inherent variation in centrifugal forces on a rotor blade to achieve optimum blade twist distributions in each flight mode through the use of extension-twist coupled composite rotor blades, has been considered for performance improvement of tiltrotor aircraft over the last two decades. The challenge for this concept is to achieve the desired twisting deformations in the rotor blade without altering the aeroelastic characteristics of the vehicle. A concept referred to as the sliding mass concept is proposed in this work in order to increase the twist change with rotor speed for a closed-cell composite rotor blade cross-section to practical levels for performance improvement in a tiltrotor aircraft. The concept is based on load path changes for the centrifugal forces by utilizing non-structural masses readily available on a conventional blade, such as the leading edge balancing mass. A multilevel optimization technique based on the simulated annealing method is applied to improve the performance of the XV15 tiltrotor aircraft. A cross-sectional analysis tool, VABS together with a multibody dynamics code, DYMORE are integrated into the optimization process. The optimization results revealed significant improvements in the power requirement in hover while preserving cruise efficiency. It is also shown that about 21% of the improvement is provided through the sliding mass concept pointing to the additional flexibility the concept provides for tailoring of the structure without any additional weight penalty on the system.

Offshore wind power production is an attractive clean energy option, but the difficulty of access can lead to expensive and rare opportunities for maintenance. Smart loads management (controls) are investigated for their potential to increase the fatigue life of damaged offshore wind turbine rotor blades. This study will consider two commonly encountered damage types for wind turbine blades, the trailing edge disbond (bond line failure) and shear web disbond, and show how 3D finite element modeling can be used to quantify the effect of operations and control strategies designed to extend the fatigue life of damaged blades. Modern wind turbine blades are advanced composite structures, and blade optimization problems can be complex with many structural design variables and a wide variety of aeroelastic design requirements. The multi-level design method is an aeroelastic structural design technique for beam-like structures in which the general design problem is divided into a 1D beam optimization and a 2D section optimization. As a demonstration of aeroelastic design, the multi-level design method is demonstrated for the internal structural design of a modern composite rotor blade. Aeroelastic design involves optimization of system geometry features as well as internal features, and this is demonstrated in the design of a flying wing aircraft. Control methods such as feedback control also have the capability alleviate aeroelastic design requirements and this is also demonstrated in the flying wing aircraft example. In the case of damaged wind turbine blades, load mitigation control strategies have the potential to mitigate the effects of damage, and allow partial operation to avoid shutdown. The load mitigation strategies will be demonstrated for a representative state-of-the-art wind turbine (126m rotor diameter). An economic incentive will be provided for the proposed operations strategies, in terms of weighing the cost and risk of implementation against the benefits of increased revenue due to operation of damaged turbines. The industry trend in wind turbine design is moving towards very large blades, causing the basic design criterion to change as aeroelastic effects become more important. An ongoing 100 m blade (205 m rotor diameter) design effort intends to investigate these design challenges. As a part of that effort, this thesis will investigate damage tolerant design strategies to ensure next-generation blades are more reliable.

The content of study is related to reducing the production times of milling of aluminum component of vacuum pump. The primary aim of the thesis is to experimentally verify the proposed cutting conditions when milling the rotor blades and then evaluate them according to predetermined parameters. The output is reducing machining time with respect to cost, tools and machine.

Helicopter aeroelastic analysis is highly complex and multidisciplinary in nature; the flexibility of main rotor blades is coupled with aerodynamics, dynamics and control systems. A key component of an aeroelastic analysis is the vehicle trim procedure. Trim requires calculation of the main rotor and tail rotor controls and the vehicle attitude which cause the six steady forces and moments about the helicopter center of gravity to be zero. Trim simulates steady level flight of the helicopter. The trim equations are six nonlinear equations which depend on blade response and aerodynamic forcing through finite element analysis. Simulating the behavior of the helicopter in flight requires the solution of this system of nonlinear algebraic equations with unknowns being pilot controls and vehicle attitude angles. The nonlinear solution procedure is prone to slow convergence and occasional divergence causing problems in optimization and stochastic simulation studies. In this thesis, an attempt is made to efficiently solve the nonlinear equations involved in helicopter trim. Typically, nonlinear equations in mathematical physics and engineering are solved by linearizing the equations and forming various iterative procedures, then executing the numerical simulation. Helicopter aeroelasticity involves the solution of systems of nonlinear equations in a computationally expensive environment. The Newton method is typically used for the solution of these equations. Due to the expensive nature of each aeroelastic analysis iteration, Jacobian calculation at each iteration for the Newton method is not feasible for the trim problems. Thus, the Jacobian is calculated only once about the initial trim estimate and held constant thereafter. However, Jacobian modifications and updates can improve the performance of the Newton method. A comparative study is done in this thesis by incorporating different Jacobian update methods and selecting appropriate damping schemes for solving the nonlinear equations in helicopter trim. A modified Newton method with varying damping factor, Broyden rank-1 update and BFGS rank-2 update are explored using the Jacobian calculated at the initial guess. An efficient and robust approach for solving the strongly coupled nonlinear equations in helicopter trim based on the modified Newton method is developed. An appropriate initial estimate of the trim state is needed for successful helicopter trim. Typically, a guess from a simpler physical model such as a rigid blade analysis is used. However, it is interesting to study the impact of other starting points on the helicopter trim problem. In this work, an attempt is made to determine the control inputs that can have considerable effect on the convergence of trim solution in the aeroelastic analysis of helicopter rotors by investigating the basin of attraction of the nonlinear equations (set of initial guess points from which the nonlinear equations converge). It is illustrated that the three main rotor pitch controls of collective pitch, longitudinal cyclic pitch and lateral cyclic pitch have significant contribution to the convergence of the trim solution. Trajectories of the Newton iterates are shown and some ideas for accelerating the convergence of trim solution in the aeroelastic analysis of helicopter are proposed.