Abstract The interaction between the fuel jet, the oil jet and the airflow is involved in the afterburner (or ramjet combustion chamber) and the lubricating oil system of the aero-engine respectively. The latter mainly studies the penetration depth of the oil jet into the airflow, the oil jet breakup position and so on. In the under-race lubrication system, the oil jet is deflected due to the high-speed rotation of the oil scoop and some droplets, ligaments are separated. The deflection of the oil jet and the splash of droplets may affect the oil capture efficiency of the under-race lubrication system. At the same time, the configuration of the oil jet nozzle will also have a certain impact on the oil capture efficiency. Therefore, it is necessary to carry out research on the flow characteristics of the oil jet in the airflow, and provide reference for the oil jet nozzle configurations of the under-race lubrication system. In this paper, the calculation results show that the Couple Level-Set and Volume of Fluid (CLSVOF) method is better than the Volume of Fluid (VOF) method. The correlations between the coordinate of the oil jet breakup positions and the liquid-air momentum ratio were concluded. The equation of the trajectory curve was derived for the jet column trajectory before breakup. The difference of the oil jet flow characteristics between single jet nozzle and the twin jet nozzle and the tandem jet nozzle configuration is also studied. Finally, the design method under the tandem jet nozzle configuration is given.

Abstract Intershaft bearing is widely adopted in dual rotor turbofan aircraft engines. Since this kind of dual rotor system has two different rotor speeds and the intershaft bearing leads to the coupling between HP rotor and LP rotor, the calculation of the critical speeds is much more complicated than that of the rotor systems without intershaft bearing. Compared to a single rotor system, the dual rotor system has more critical speeds which can be classified as critical speeds excited by HP rotor and that by LP rotor. In the paper, a finite element rotor model of a high-bypass turbofan jet engine with intershaft bearing is established for the study of critical speeds analysis. The general axisymmetric element is used to model the shafts and disks, and the blades are simplified to mass points. The main bearings including the intershaft bearing are set up with spring element. Assuming that the rotational speed ratio of the two rotors for the dual rotor system is a fixed number, the critical speeds are calculated using three methods based on the finite element rotor model. For the first method, the system critical speeds are obtained directly by Campbell diagram based on QR damped solution method. Then the synchronous unbalance response analyses are carried out and the rotor critical speeds are derived from the amplitude-frequency curves. For the last method, multiple group Campbell diagram analyses are conducted. With one rotor speed fixed at constant rpm N, we can change the speed of the other rotor to obtain one group of critical speeds. By varying speed N of the two rotors, a critical speeds data set can be obtained and plotted as a dual rotor critical speed map. The critical speeds can be easily extracted from the critical speed map according to the rotational speed curve of the engine. The study shows that the dual rotor system critical speeds calculated from above three methods are identical. For the first two methods, the rotational speed ratio of two rotors must be a known and fixed number, which is impossible in reality. The third proposal has no rotation speed relation restriction for rotors, and therefore is recommended for analyzing the critical speeds of aircraft engines with intershaft bearing.

Abstract Contra-rotation has several advantages like swirl-free discharge, high pressure-rise per stage, and possibility of operating both the rotors at different speeds. With these merits, contra-rotating fan emerges as a competent technology for future gas turbine engines. During operation, gas turbine engines undergo situations like high angle of attack manoeuvres, large cross-winds, bird-hits, etc. which distort the flow at the inlet of the engine. A thorough understanding of the effect of distortion on low aspect ratio contra-rotating fans is missing in literature. This paper reports the consequences of radial distortion on the performance of a low aspect ratio contra-rotating fan. The uniform inlet flow is distorted radially using wire mesh screens. The unsteady data obtained from high response pressure sensors are analysed using Discrete Spatial Fourier Series (DSFS) and Morlet wavelet transform. Both Long Length Scale Disturbances (LLSD) or modal waves and Short Length Scale Disturbances (SLSD) or spikes are observed for different inflow conditions. The stage stalls primarily due to the instabilities arising at the tip region of rotor-1. Rotor-2 shows poor coherence in the disturbances prior to stall compared to that of rotor-1. Tip-distorted flow is dominated with SLSDs in the pre-stall region and hence a stall precursor is not observed whereas clean and hub-distorted flows show prominent LLSDs prior to stall. The radial distortions get redistributed at the exit of rotor-1 and hence, the distorted inflows do not severely lead to instabilities on rotor-2. In summary, this work explains in detail the development of unsteady flow phenomena occurring in a low-aspect ratio contra-rotating fan stage leading to stall and the way in which the system responds to it.

Abstract Experimental investigation on the plasma flow control of axial compressor rotating stall is implemented in this paper. The control effects of axial plasma actuation (inducing body force of compressor axial direction) with three different locations as well as stagger angle plasma actuation (inducing body force perpendicular to the compressor rotor tip chordwise direction) at different rotation speeds are studied. An unsteady plasma actuation is designed to influence the unsteady rotor tip flow at near stall point, and is found to be the most powerful in improving the compressor stall margin. Both the compressor rotation speed and the plasma actuation voltage are found to be very influential on the control effects of the plasma actuation. The abilities of the plasma actuation in suppressing the compressor rotating stall and influencing the compressor static pressure rise coefficient are not directly related. All the plasma actuations studied can improve the compressor stall margin, but the compressor static pressure rise coefficient can be decreased or increased with different plasma actuation layouts.

Abstract Performance and flow measurements are carried out to investigate the impact of varying the geometry of axial casing grooves on the stall margin and efficiency of an axial turbomachine. Prior studies have shown that skewed semi-circular grooves installed near the blade leading edge (LE) have multiple effects on the flow structure, including ingestion of the tip leakage vortex (TLV), suppression of backflow vortices, and periodic variations of flow angle. To determine which of these phenomena is a key contributor, the present study examines the impact of several grooves, all with the same inlet geometry, but with outlets aimed at different directions. The “U” grooves that have circumferential exits aimed against the direction of blade rotation achieve the highest stall margin improvement of well above 60% but cause a 2.0% efficiency loss near the best efficiency point (BEP). The “S” grooves, which have exits aimed with the blade rotation, achieve a relatively moderate stall margin improvement of 36%, but they do not reduce the BEP efficiency. Other grooves, which are aligned with and against the flow direction at the exit from upstream inlet guide vanes, achieve lower improvements. These trends suggest that causing high periodic variations in flow angle around the blade leading edge is particularly effective in extending the stall margin, but also reduces the peak efficiency. In contrast, maintaining low flow angles near the LE achieves more moderate improvement in stall margin, without the maximum efficiency loss. Hence, of the geometries tested, the S grooves appear to have the best overall impact on the machine performance. Velocity measurements and flow visualizations are performed in an axial plane located downstream of the grooves, near the trailing edge of the rotor. Reduced efficiency or performance co-occurs with elevated circumferential velocity in the tip region, but differences in the axial blockage are subtle. Yet, near the BEP, the regions with reduced axial velocity, or even negative velocity between the TLV and the endwall, are wider behind the U grooves compared to the S grooves. The vorticity profiles also show that at low flow rates the TLV is ingested entirely by the grooves, in contrast to the best efficiency point, where a considerable fraction of the TLV rollup occurs downstream of the grooves.

Abstract Counter-rotating turbomachines have the potential to be high efficiency, high power density devices. Comparisons between conventional and counter-rotating turbomachines in the literature make multiple and often contradicting conclusions about their relative performance. By adopting appropriate non-dimensional parameters, based on relative blade speed, the design space of conventional machines can be extended to include those with counter-rotation. This allows engineers familiar with conventional turbomachinery to transfer their experience to counter-rotating machines. By matching appropriate non-dimensional parameters the loss mechanisms directly affected by counter-rotation can be determined. A series of computational studies are performed to investigate the relative performance of conventional and counter-rotating turbines with the same non-dimensional design parameters. Each study targets a specific loss source, highlighting which phenomena are directly due to counter-rotation and which are solely due to blade design. The studies range from two-dimensional blade sections to three-dimensional finite radius stages. It is shown that, at hub-to-tip ratios approaching unity, with matched non-dimensional design parameters, the stage efficiency and work output are identical for both types of machine. However, a counter-rotating turbine in the study is shown to have an efficiency advantage over a conventional machine of up to 0.35 percentage points for a hub-to-tip ratio of 0.65. This is due to differences in absolute velocity producing different spanwise blade designs.

Abstract The efficiency of modern axial turbomachinery is strongly driven by the secondary flows within the vane or blade passages. The secondary flows are characterized by a complex pattern of vortical structures that origin, interact and dissipate along the turbine gas path. The endwall flows are responsible for the generation of a significant part of the overall turbine loss because of the dissipation of secondary kinetic energy and mixing-out of non-uniform momentum flows. The understanding and analysis of secondary flows requires a reliable vortex identification technique to predict and analyse the impact of specific turbine designs on the turbine performance. However, literature shows a remarkable lack of general methods to detect vortices and to determine the location of their cores and to quantify their strength. This paper presents a novel technique for the identification of vortical structures in a general 3D flow field. The method operates on the local flow field and it is based on a triple decomposition of motion proposed by Kolář. In contrast to a decomposition of velocity gradient into the strain and vorticity tensors, this method considers a third, pure shear component. The subtraction of the pure shear tensor from the velocity gradient remedies the inherent flaw of vorticity-based techniques which cannot distinguish between rigid rotation and shear. The triple decomposition of motion serves to obtain a 3D field of residual vorticity whose magnitude is used to define vortex regions. The present method allows to locate automatically the core of each vortex, quantify its strength and determine the vortex bounding surface. The output may be used to visualize the turbine vortical structures for the purpose of interpreting the complex three-dimensional viscous flow field, as well as to highlight any case-to-case variations by quantifying the vortex strength and location. The vortex identification method is applied to a high-pressure turbine with three optimized blade tip geometries. The 3D flow-field is obtained by CFD computations performed with Numeca FINE/Open. The computational model uses steady-state RANS equations closed by the Spalart-Allmaras turbulence model. Although developed for turbomachinery applications, the vortex identification method proposed in this work is of general applicability to any three-dimensional flow-field.

Abstract Non-conventional diffuser designs are introduced to minimize the energy losses associated with diffusion and to enhance stable operating range of the diffusion system. This is achieved by reducing width of the diffuser by modifying a hub or shroud curves by keeping the diffuser diameter ratio constant which is often known as pinch. The comparison of modified compressor with base model is accomplished by using performance characteristics such as static pressure recovery coefficient, stagnation pressure loss coefficient, energy coefficient and isentropic efficiency. Simulations are performed at various hub pinch (5%–20%), shroud pinch (5%–20%) and combined hub and shroud pinch (5%–20%). Among all the cases shroud pinch of 10% shows best results in terms of stagnation pressure loss coefficient, static pressure recovery coefficient and energy coefficient. Further, simulations are carried out with forced rotating vaneless diffuser. It gives better results in terms of pressure rise with lower stagnation pressure losses. But there is a moderate decrement in isentropic efficiency of compressor when compared to the base model.

Abstract In aerodynamic design, accurate and robust surrogate models are important to accelerate computationally expensive CFD-based optimization. Machine learning techniques can also enable affordable exploration of high-dimensional design spaces with targeted selection of sparse high-fidelity data. In this paper, a multi-fidelity global-local approach is presented and applied to the surrogate-based design optimization of a highly-loaded transonic compressor rotor. The key idea is to train multi-fidelity surrogates with fewer high-fidelity RANS predictions and more rapid and inexpensive lower-fidelity RANS evaluations. The framework also introduces a global-local search algorithm that can spin-off multiple local optimization threads over narrow and targeted design spaces, concurrently to a constantly adapting global optimization thread. The approach is demonstrated with an optimization of the transonic NASA rotor 37, yielding significant increase in performance within a dozen of optimization iterations.

Abstract This paper focuses on rotor dynamic investigation of a bladeless turbine, or Tesla turbine, for application to innovative small scale cycles. Tesla rotor consists of a shaft with several co-rotating disks with small gaps between each other. The flow through the disks creates a momentum exchange by viscous effect, motoring the shaft. Thanks to its simplicity and low cost, the Tesla expander is attractive for energy harvesting and waste heat recovery from low/medium temperature in small and micro scale applications. Rotor assembly and its parts may present dynamic criticalities, due to their structural characteristics: to predict and ensure low vibrations during operations, numerical and experimental studies have been carried out on some prototypes. The activity started considering a non-rotating single Tesla disk both in free and real constrained configuration: an experimental modal analysis was performed, whose results were used to validate a disk numerical model. In this case, an analytical approach with a simplified geometry assumption was considered. All methods results were correlated each other and discrepancies have been identified and analysed. Furthermore, the investigation of a single disk rotor vibrational behaviour has been extended from static conditions to rotating conditions. Numerical analysis has been carried on taking into account the effect of gyroscopic couples and centrifugal field generated by disk rotation. In parallel, a corresponding experimental activity has been done using a dedicated test rig which allowed to perform vibrational operational measurement while the disk was in motion. Campbell diagram of the single rotating disk on the shaft has been obtained from numerical and experimental analysis allowing to identify system dynamic behaviour and to deepen aspects related to critical speeds. Finally, a whole rotor model has been developed, allowing the characterization of the dynamic behaviour of a fully assembled turbine rotor. The developed models, validated with experiments, are powerful tools that can predict the bladeless expander vibrational behaviour at the early phase of design.

Abstract Combustion characteristics of a can combustor with a rotating casing for an innovative micro gas turbine have been modeled. The effects of syngas compositions and the rotating speed on the combustor performance were investigated. The effects of rotation on the combustion performance have been studied previously with methane as the fuel. This work extended the investigation for future application with syngas blended fuels. Two typical compositions of syngas were used namely: H 2 -rich (H 2 :CO=80:20, by volume) and equal molar (H 2 :CO=50:50). The analyses were performed with a computational model, which consists of three-dimension compressible k-ε realizable turbulent flow model and presumed probability density function for combustion process invoking a laminar flamelet assumption generated by detailed chemical kinetics from GRI 3.0. As syngas is substituted for methane at a constant fuel flow rate, the high temperature flame is stabilized along the wall of the combustor liner. With the casing rotating, pattern factor and exit temperature increase, but the lower heating value of syngas causes a power shortage. To make up the power, the fuel flow rate is raised to maintain the thermal load. Consequently, the high temperature flame is pushed downstream due to increased fuel injection velocity. NO x emission decreases as the rotational speed increases in both cases. Pattern factor decreases but exit temperature increases with the increase of roatation speed indicating a higher combustion efficiency. However, there is possible hotspots at exit due to higher pattern factor (PF>0.3) for H 2 -rich and equal molar syngas at lower speed of rotation, which needs to be resolved by improving the cooling strategy.

Abstract This work presents the results of a numerical analysis performed on a gas turbine leading edge cooling system. The investigation was carried out in order to provide a detailed interpretation of the outcomes of a parallel experimental campaign. The cooling geometry consists of a cold bridge type impingement system: a radial channel feeds an array of holes, which in turn generate impingement jets cooling down the inner side of the leading edge surface. Coolant is extracted by five rows of holes, replicating film cooling and showerhead systems. Two impingement geometries were considered, presenting different holes arrangements and diameters but sharing the same overall passage area, in order to highlight the effect of different coolant distributions inside the leading edge cavity. For both geometries a single test point was investigated in static and rotating conditions, with an equivalent slot Reynolds number of around 8200 and feeding conditions corresponding to the midspan radial section of the blade. Both steady RANS approach and Scale Adaptive Simulation (SAS) were tested. Due to the strong unsteadiness of the flow field, the latter proved to be superior: as a consequence, SAS approach was adopted to study every case. A fairly good agreement was observed between the measured and computed heat transfer distributions, which allowed to exploit the numerical results to get a detailed description of the phenomena associated with the different cases. Results reveal that the two holes arrangements lead to strongly different heat transfer patterns, related to the specific flow phenomena occurring inside the leading edge cavity and to the mutual influence of the various system features. Rotational effects also appear to interact with the supply condition, altering the jet lateral spreading and the overall heat transfer performance.

Abstract High thermal load on the turbine blade tip surface leads to high temperature corrosion and severe structural damage. One common way is to deliver a part of coolant through bleed holes onto the tip portion for cooling purpose. In this study, numerical simulations are performed to investigate the effects of rotation on the internal tip heat transfer in a simplified rotating two-pass channel with a bleed hole, which is applicable to the internal cooling passage of typical gas turbine blade. The bleed hole is placed on the tip wall of a two-pass channel at different locations, i.e. the ratio of distance from the outlet-side wall to width of the tip wall is 0.07, 0.21, 0.5, 0.78, 0.93, respectively. A smooth channel without bleed hole is used as Baseline. The Reynolds number is fixed at 10,000. The Ro numbers are varied from 0 to 0.4. Results show that a three-dimensional vortex, which is induced by the Coriolis force, is found at the bend region. It transports the fluid from the trailing side to leading side, which is beneficial to enhance tip heat transfer. The middle-mounted hole shows a better heat transfer augmentation compared to other hole arrangement. The rotation have a notable effect on the heat transfer and flow structures. Compared to the smooth channel, the heat transfer augmentation is about 34%.

Abstract The double turning areas ribbed serpentine channel with lateral outflow is an important structure for designing the internal systems of turbine blade. The current work similarly simplifies the internal channel of the real blade. The Nusselt number and pressure coefficient distribution of the double turning areas ribbed serpentine channel with different outflow ratios are numerically researched under static and rotating conditions. The Realizable k-ε turbulence model with enhanced wall treatment is used in the numerical simulation. The inlet Reynolds number is 11000. The rotation numbers vary from 0 to 0.09. Three outflow ratios are 27%/0%/73%, 27%/49%/24% and 27%/73%/0%, respectively. The rotation radius (R) is 46.4d. The result shows that the Nusselt number distribution of the passage 3 under 27%/49%/24% outflow ratio condition is similar to that under 27%/73%/0% outflow ratio condition. There is a large low Nusselt number area in the passage 3 under Dr = 27%/0%/73% condition. The averaged area Nusselt number ratios on the suction side of the passage 1, passage 2 and passage 3 are higher than that on the pressure side under nonrotating condition. Rotation enhances heat transfer on the suction side of the passage 2, and has a positive effect on pressure side heat transfer of passage 1 and passage 3. The averaged area Nusselt number ratio of passage 3 under 27%/73%/0% outflow ratio condition is higher than that under other outflow ratio conditions. With the rotation number increasing, the pressure coefficient of the complete ribbed serpentine channel gradually increases, and the maximum increase is in the first turning area.

Abstract In the present work, the turbulent flow fields in a static and rotating ribbed channel representative of an aeronautical gas turbine are investigated by mean of wall-resolved compressible Large Eddy Simulation (LES). This approach has been previously validated in a squared ribbed channel based on an experimental database from Von Karman Institute (Reynolds and rotation numbers of about 15000 and +/− 0.38 respectively). LES results prove to reproduce differences induced by buoyancy in the near rib region and resulting from adiabatic or anisothermal flows under rotation. The model also manages to predict the turbulence increase (decrease) around the rib in destabilizing (stabilizing) rotation of the ribbed channels. On this basis, this paper investigates in more details the spatial development of the flow along the channel and its potential impact on secondary flow structures. More specifically and for all simulations, results on the adiabatic static case exhibits two contra-rotating structures close to the lateral walls of the channel induced by transversal pressure difference created by the ribs. These structures are generated after the first ribs and appear behind all inter rib sections, their relative position being partly affected by rotation. When considering the stabilizing rotating case, two additional contra-rotating structures also develop along the channel from the entrance close to the low-pressure wall (rib-mounted side). These vortices are due to the confinement of the configuration, inflow profile and are the result of Coriolis forces induced by rotation. Görtler vortices also appear on the pressure wall (opposite to the rib-mounted side). In the destabilizing rotating case, these two types of secondary structures are found to co-exist and their migration in the channel is significantly different due to the presence of the ribs on the pressure side. Finally, it is shown that heat transfer affects only marginally the static and stabilized cases while it changes more significantly the flow organization in the destabilizing case mainly because of enhanced heat transfer and increased buoyancy force effects.

Abstract The internal cooling passages of gas turbine blades mostly have varying aspect ratios from one passage to another. However, there are limited data available in the open literature that used a reduced cross-section and aspect ratio, AR, after the tip turn. Therefore, the current study presents heat transfer and pressure drop of three different α = 45° profiled rib orientations, typical parallel (usual), reversed parallel (unusual), and criss-cross patterns in a rotating two-pass rectangular channel with AR = 4:1 and 2:1 in the first radially outward flow and second radially inward flow passages respectively. For each rib orientation, regional averaged heat transfer results are obtained for both the flow passages with the Reynolds number ranging from 10,000 to 70,000 for the first passage and 16000 to 114000 for the second passage with a rotational speed range of 0 rpm to 400 rpm. This results in the highest rotation number of 0.39 and 0.16 for the first and second passage respectively. The effects of rib orientation, aspect ratio variation, 180° tip turn, and rotation number on the heat transfer and pressure drop will be addressed. According to the results, for usual, unusual and criss-cross rib patterns, increasing rotation number causes the heat transfer to decrease on the leading surface and increase on the trailing surface for the first passage and vice versa for the second passage. Overall heat transfer enhancement of the usual and unusual rib patterns is higher than criss-cross one. In terms of the pressure losses, the criss-cross rib pattern has the lowest and the usual rib pattern has the highest-pressure loss coefficients. When pressure loss and heat transfer enhancement are both taken into account together, the criss-cross or unusual rib pattern might be an option to use in the internal cooling method. Therefore, the results can be useful for turbine blade internal cooling design and heat transfer analysis.

Abstract In the current work, the effect of Coriolis force is considered in a rotating rectangular rib-roughened channel. Time-resolved PIV (Particle Image Velocimetry) was used to measure the flow field. The Reynolds number is fixed at 10000, and the rotation numbers range from 0 to 0.52. The ribs obstruct the channel by 10%. The mean velocity fields and Reynolds shear stress fields are obtained. Due to the effect of Coriolis force, the flow fields are different between leading and trailing side. Furthermore, the vortex and the reattachment are also investigated. Based on the results, it can be inferred that Coriolis force plays a significant effect on the vortices. Coriolis force enlarges the vortex near the leading side, but suppresses the vortex near the trailing side. An interesting phenomenon has been found. Because of the Coriolis force pointing to the trailing side, the main stream is supposed to be pushed to the trailing side. But the results show that a velocity deficit appears near the trailing side. This phenomenon indicates that the Coriolis force is not the only force affecting the flow, but the secondary flow is also an important factor that cannot be ignored in a rotating ribbed channel.

Abstract The current study experimentally studied heat transfer characteristics of the blade tip wall in a rotating internal cooling channel. The aspect ratio of this rectangular channel was 1:4, and the hydraulic diameter was 25.6 mm. Due to the impact of the 180° turn, complex three dimensional flow significantly affected heat transfer on the internal tip surface. The liquid crystal method is used to capture the heat transfer contour on the internal tip surface. In this study, the leading and trailing surfaces of the channel wall were either smooth or roughened with 45° angled ribs. The Reynolds number inside the pressurize two-pass cooling channel ranged from 10,000 to 30,000 at both stationary and rotating conditions. Furthermore, two channel orientations (90° and 135°) were tested. The effect of Coriolis force on heat transfer is studied with the rotation number up to 0.53. The tip heat transfer from the smooth channel wall was more sensitive to rotation and the largest heat transfer enhancement as a result of rotation was 68%.

Abstract This paper presents a complex approach to designing ceramic blades dovetail joints. Two ceramic materials were considered: diamond reinforced silicon carbide and hot pressed silicon nitride (only in contact testing). The model blades were made from diamond reinforced silicon carbide due to its availability and mature technology. Part 1 describes the models of mechanical and thermal contact and contains experimental data on contact strength and thermal contact conductance. Part 2 investigates the effect of stress concentration and scale factor on the fracture of ceramic test pieces. Part 3 proposes a theoretical-experimental method to estimate friction coefficient in a dovetail joint. We also investigated the character of the ceramic blades fracture during the rotor spin-up. Part 4 deals with automating the design process of the ceramic blades dovetail joints using experimental data.

Abstract Additive manufacturing (AM), also known as 3D printing, is a rapidly developing technology with tremendous potential in both developmental and production applications. Solar Turbines Incorporated is committed to AM technology for gas turbine applications. The ability to metal 3D print novel designs of turbine blades capable of actual turbine engine operation would effectively reduce design validation cycle time, and allow acquisition of key performance data early in a design campaign. In support of Solar’s advanced manufacturing development and ongoing engine efficiency improvement goals, Solar initiated a project to print a full set of Mercury™ 50 stage 2 turbine blades to be run in a development engine. Solar leveraged years of experience with design and serial production of AM components in support of this project. A significant challenge faced when printing turbine blades is producing metal with mechanical properties sufficient to withstand the rigors of engine operation. As a rotating component within the hot section of the engine, turbine blades experience high centrifugal and pressure loads at elevated temperatures. After investigation of possible alloys capable of meeting the requirements of the Mercury™ 50 design envelope, the gamma prime (γ’) strengthened nickel superalloy Inconel™ 738LC was selected to provide the best opportunity for successful development engine testing. Solar partnered with Oak Ridge National Laboratory (ORNL) to produce the Inconel™ 738LC blades with Electron Beam Melting (EBM) powder bed fusion process. Once a rough blade shape was printed, the fir-tree attachment, blade tip shroud, and air flow path surfaces were finished using both conventional and non-conventional machining processes. In-process monitoring, metallurgical evaluations, mechanical testing, and non-destructive inspection techniques were used to validate the printed blade material integrity and conformance to geometric design intent. Planned future activities include assembly of the AM blades onto a disk for spin pit testing to validate the mechanical integrity and design margin of the blades. The final phase of the project will be to install the bladed disk assembly into a Mercury™ 50 engine at Solar Turbines to conduct a series of hot-fired engine performance tests.