Abstract A methodology is presented for the re-design of a large centrifugal impeller used for the lift fan on an air cushion vehicle. The design is driven by stringent requirements for aerodynamic and structural performance. It is also desired to minimize the fan’s life cycle cost, by reducing both the manufacturing cost and on-going maintenance burden. Improving the fan efficiency will increase the vehicle’s endurance and range, and minimizing life cycle cost will reduce the overall operational expenses. The lift fan assembly has a double-inlet impeller and an offset double-discharge volute. The new impeller design provides increased air flow with similar aerodynamic efficiency when compared to the prior design. To reduce the manufacturing cost, the new fan blade design can be produced by an aluminum extrusion process. The manufacturing process dictates that the blade cross-section be two-dimensional across the entire span, which poses structural challenges for attachment of the blades to the shroud and center disk. Analysis of the aerodynamic and structural performance of the lift fan was carried out using computational fluid dynamics (CFD) and structural finite element analysis (FEA) models. The fabrication of full-scale lift fan components is described. Plans for final assembly and for conducting full-scale, full-power aerodynamic testing will also be explained.

Abstract In the last decades major improvements in transonic centrifugal compressor design have been achieved. The further exploration of design space is enabled by recent progress in structural mechanics and manufacturing. A challenging task of inducer design especially in terms of transonic inflow conditions is to provide a wide flow range and reduced losses due to a sufficient shock control. The use of so called multidisciplinary design optimization with an extensive amount of free parameters leads finally to complex designs. DLR’s latest Fast Rotating Centrifugal Compressor (SRV5) operates at a design speed of M u2 = 1.72 and a total pressure ratio of 5.72. This compressor design is characterized by an S-shaped leading edge and free-form blade surfaces. Due to the complex design the key design features are difficult to explore. Therefore, non-intrusive measurements are conducted on the highly loaded SRV5. The Laser-2-Focus (L2F) approach that is used in addition with the Doppler Global Velocimetry (DGV) delivers a three dimensional velocity field. Besides the impeller inflow the ouflow is also part of the experimental and numerical verification of the advanced compressor design. Experimental results are compared with the numerical analysis of the compressor using DLR’s RANS Flow Solver TRACE. The deep insight of the inflow leads to a better understanding of the operating behavior of such impeller designs.

Abstract An additive manufactured (AM) vaned diffuser for use in a centrifugal compressor research facility was designed and implemented. Utilizing an AM process to manufacture the diffuser reduces the long lead time that is associated with conventionally manufactured diffusers, and it increases the instrumentation capabilities within the flow path. Several AM techniques and a variety of plastic and metal materials were evaluated for this application. A high-temperature, stereolithography (SL) resin was chosen because of the tight dimensional tolerances maintained by the SL process. Utilizing a high-temperature plastic also results in manufacturing costs that are significantly less than using a metal material. Samples of the chosen material were subjected to mechanical testing to investigate the effects of build direction (BD) and to verify its properties in the high-temperature compressor environment. To fit within the manufacturing space of an SL machine, the AM diffuser consists of seven radially symmetric sections that are assembled to form a complete flow path. Considerations for modifying the research facility to allow for this unique installation are presented. Precision measurements of the AM components were obtained to compare printed and modeled geometry, and they demonstrate close alignment of flow path dimensions.

Abstract This paper presents a multidisciplinary adjoint-based design optimization of a turbocharger radial turbine for automotive applications. The aim is to improve the total-to-static efficiency of the turbine while keeping mechanical stresses below a predefined limit. The search for the optimal design is accomplished using an efficient Sequential Quadratic Programming algorithm considering additional aerodynamic and manufacturing constraints. The aerodynamic performance of the wheel is evaluated by a Reynolds-Averaged Navier-Stokes solver, whereas the maximum stresses in the material are predicted by a Finite Element Analysis tool. The design gradients required by the optimizer are computed with the adjoint approach which provides sensitivity information largely independent of the number of design variables. The results presented in this paper show the clear need to take into account mechanical stresses during optimization, as they are the most restrictive design limitation. However, the gradient-based optimization algorithm is able to effectively keep the stress levels below the critical value while significantly improving the turbine efficiency in a few design cycles.

Abstract Turbomachinery blade design is based on an axisymmetric model assuming identical blades in the circumferential direction. The blade manufacturing process, however, has a finite accuracy leading to surface imperfections and therefore non-identical blades. This leads to a variation in each passage cross area effecting the aerodynamic performance of the turbine. The simulation of multiple 3D CFD blade passages, nevertheless, is still computational expensive. However, if the effect of nonaxisymmetric blades is small enough to be approximated as linear, the adjoint method is an inexpensive alternative, which can be used in an industrial context. A set of 102 turbine vanes is used to model the manufacturing variations and assess the aerodynamic impact of up to eight nonaxisymmetric blade passages of a transonic high pressure turbine stage. It is shown that the modeled blade passage results superimpose linearly. Therefore, the aerodynamic impact of surface imperfections can be evaluated by superpositioning the singlepassage results, which are generated by an adjoint computation.

Abstract Nowadays a preliminary assessment on environmental impact of a new product is becoming more and more important. It is useful for a designer to access to a comprehensive methodology that supports configuration assessments taking into account the whole product lifecycle from the beginning of conceptual phase. To develop a competitive product, and particularly a gas turbine, each design trade-off needs to be performed considering not only the typical parameters such as performances, life and costs but also the cradle-to-grave environmental impact. Scope of the following paper is the application of design-for-Environment methodology to different architectures of GT compressor rotor module. Three design alternatives are analyzed and compared in terms of ELCA considering their design, material selection, manufacturing process and operating life. Specific considerations are proposed as a result of the combination of traditional design practices with environmental assessment. This study highlighted that number of parts, weight and amount of material removed or scraped that is, in other words, the level of production process optimization, are the key factors to control the environmental impact of a product.

Abstract Thermal barrier coatings (TBC) found on turbine blades are a key element in the performance and reliability of modern gas turbines. TBC reduces the heat transfer into turbine blades by introducing an additional surface thermal resistance; consequently allowing for higher gas temperatures. During the service life of the blades, the TBC surface may be damaged due to manufacturing imperfections, handling damage, service spalling, or service impact damage, producing chips in the coating. While an increase in aerofoil temperature is expected, it is unknown to what degree the blade will be affected and what parameters of the chip shape affect this result. During routine inspections, the severity of the chipping will often fall to the discretion of the inspecting engineer. Without a quantitative understanding of the flow and heat transfer around these chips, there is potential for premature removal or possible blade failure if left to operate. The goal of this preliminary study is to identify the major driving parameters that lead to the increase in metal temperature when TBC is damaged, such that more quantitative estimates of blade life and refurbishing needs can be made. A two-dimensional computational Conjugate Heat Transfer model was developed; fully resolving the hot gas path and TBC, bond-coat, and super alloy solids. Representative convective conditions were applied to the cold side to emulate the characteristics of a cooled turbine blade. The hot gas path properties included an inlet temperature of 1600 K with varying Mach numbers of 0.30, 0.59, and 0.80 and Reynolds number of 5.1×10 5 , 7.0×10 5 , and 9.0×10 5 as referenced from the leading edge of the model. The cold side was given a coolant temperature of 750 K and a heat transfer coefficient of 1500 W/m 2 *K. The assigned thermal conductivities of the TBC, bond-coat, and metal alloys were 0.7 W/m*K, 7.0 W/m*K, and 11.0 W/m*K, respectively, and layer thicknesses of 0.50 mm, 0.25 mm, and 1.50 mm, respectively. A flat plate model without the presence of the chip was first evaluated to provide a basis of validation by comparison to existing correlations. Comparing heat transfer coefficients, the flat plate model matched within uncertainty to the Chilton-Colburn analogy. In addition, flat plate results captured the boundary layer thickness when compared with Prandtl’s 1/7th power-law. A chip was then introduced into the model, varying the chip width and the edge geometry. The most sensitive driving parameters were identified to be the chip width and Mach number. In cases where the chip width reached 16 times the TBC thickness, temperatures increased by almost 30% when compared to the undamaged equivalents. Additionally, increasing the Mach number of the incoming flow also increased metal temperatures. While the Reynolds number based on the leading edge of the model was deemed negligible, the Reynolds number based on the chip width was found to have a noticeable impact on the blade temperature. In conclusion, this study found that chip edge geometry was a negligible factor, while the Mach number, chip width, and Reynolds number based on the chip width had a significant effect on the total metal temperature.

Abstract Nowadays a preliminary evaluation of environmental impact of a new product becomes more and more important, especially when the case study refers to an industrial gas turbine both for power generation and mechanical drive applications. The environmental impact evaluation, as well as the preliminary lifecycle cost analysis, will represent a critical driver to develop a competitive product during the conceptual design phase where the engine architecture is an outcome of different alternatives trade-offs. Scope of the following paper is the presentation of a set of Design-for-Environment considerations obtained through gas turbine functional decomposition in modules, identification of the most critical, assessment of their contribution compared to the whole engine in terms of environmental impact as well as the effect on the engine use depending on ambient and operating conditions. The outcome of this study is an approach to preliminarily evaluate the engine life-cycle impact as well as a set of indications to drive machine architecture, material selection and production processes towards the sustainability during manufacturing and operational phases.

Abstract Casting deviations introduce geometric variability that impacts the aerodynamic performance of turbomachinery. These effects are studied for a High Pressure Turbine (HPT) rotor blade from a modern aero-engine. 197 blades were measured using three-dimensional structured-light scanning (GOM scanning), and the performance of each blade is quantified using Reynolds-Averaged Navier-Stokes (RANS) simulations. Casting variation is typically managed by applying geometric tolerances to determine the suitability of a component for service. The analysis demonstrates that this approach may not be optimal since it does not necessarily align with performance, in particular the capacity and efficiency. Alternatively, functional acceptance based on the predicted performance of each blade removes the uncertainty associated with geometric tolerancing and gives better performance control. Building on these findings, the paper proposes a method to set the orientation of the fir-tree, which is machined after casting. By customizing the alignment of each blade, performance variability and scrap rates can be significantly reduced. The method uses predictions of performance to reorient the castings to compensate for the manufacturing-induced errors, without changing the design-intent blade geometry and with minimal changes to the manufacturing facility.

Abstract Since the advent of premix combustion technology in industrial gas turbines, regular manual combustion tuning and engine adjustments have been necessary to maintain engines within emission regulatory limits and to control combustion dynamics (pulsations) for hardware integrity. The emissions and pulsation signatures of premix combustors are strongly driven by ambient conditions, engine performance, degradation and fuel composition. As emissions limits became more stringent over the years, higher combustion dynamics were encountered and challenges to maintain acceptable settings after yearly combustion inspections were regularly encountered. This challenge was further increased as sites operating advanced Gas Turbines (GT) eliminated Combustion Inspections (CI) and required uninterrupted generation at optimum settings for up to three years. The case for automated tuning systems became evident for the Industrial Gas Turbine (IGT) market in the mid 2000’s and different IGT manufacturers and service providers began developing them. Power Systems Manufacturing (PSM) developed the AutoTune (AT) system in 2008 and has since installed it in over fifty units, accumulating close to a million hours of operation. The history of PSM’s AT system development as well as a description of its fundamental principles and capabilities are discussed. The power generation market is changing rapidly with the injection of renewables, thus driving the demand for operational flexibility, the design of PSM’s multi-platform compatible, AutoTune system; allowing for increased peak power, extended turndown and transient tuning is discussed. The paper also describes, how, using the same tuning principles, the application for an AutoTune system can be extended to the Balance Of Plant (BOP) equipment.

Abstract Interest in small scale turbines is growing mainly for small scale power generation and energy harvesting. Conventional bladed turbines impose manufacturing limitations and higher cost, which hinder their implementation at small scale. This paper focuses on experimental and numerical performance investigation of Tesla type turbines for micro power generation. A flexible test rig for Tesla turbine fed with air as working fluid has been developed, of about 100W net mechanical power, with modular design of two convergent-divergent nozzles to get subsonic as well as supersonic flow at the exit. Seals are incorporated at the end disks to minimize leakage flow. Extensive experiments are done by varying design parameters such as disk thickness, gap between disks, radius ratio and outlet area of exhaust with speeds ranging from 10000 rpm to 40000 rpm. A quasi-1D model of the whole setup is created and tuned with experimental data to capture the overall performance. Major losses, ventilation losses at end disks, nozzle and exhaust losses are evaluated experimentally and numerically. Effect of design parameters on the performance of Tesla turbines is discussed.

Abstract Modern high pressure turbine (HPT) blade design stands out due to its high complexity comprising three-dimensional blade features, multi-passage cooling system (MPCS) and film cooling to allow for progressive thermodynamic process parameters. During the last decade, probabilistic design approaches have become increasingly important in turbomachinery to incorporate uncertainties such as geometric variations caused by manufacturing scatter and deterioration. Within this scope, the first part of this two-part paper introduces parametric models for cooled turbine blades that enable probabilistic FE analysis taking geometric variability into account to aim at sensitivity and robustness evaluation. The statistical database is represented by a population of more than 400 blades whose external geometry is captured by optical measurement techniques and 34 blades that are digitized by computed tomography (CT) to record the internal geometry and the associated variability, respectively. Based on this data, parametric models for airfoil, profiled endwall (PEW), wedge surface (WSF) and MPCS are presented. The parametric airfoil model which is based on traditional profile theory is briefly described. In this regard, a methodology is presented that enables to adapt this airfoil model to a given population of blades by means of Monte-Carlo based optimization. The endwall variability of hub and shroud are parametrized by radial offsets that are applied to the respective median endwall geometry. WSFs are analytically represented by planes. Variations of the MPCS are quantified based on the radial distribution of cooling passage centroids. Thus, an individual MPCS can be replicated by applying adapted displacement functions to the core passage centroids. For each feature that is considered within the present study, the accuracy of the parametric model is discussed with respect to the variability that is present in the investigated blade population and the measurement uncertainty. Within the scope of the second part of this paper (cf. Högner et al. [1]), the parametric models are used for a comprehensive statistical analysis to reveal the parameter correlation structure and probability density functions (PDFs). This is required for the subsequent probabilistic finite element analysis involving real geometry effects.

Abstract Modern high pressure turbine (HPT) blade design stands out due to high complexity comprising three-dimensional blade features, multi-passage cooling system (MPCS) and film cooling to allow for progressive thermodynamic process parameters. During the last decade, probabilistic design approaches have become increasingly important in turbomachinery to incorporate uncertainties such as geometric variations caused by manufacturing scatter. In part B of this two-part paper, real geometry effects are considered within a probabilistic finite element (FE) analysis that aims at sensitivity evaluation. The knowledge about the geometric variability is derived based on a blade population of more than 400 individuals by means of parametric models that are introduced in part A (cf. Högner et al. [1]). The HPT blade population is statistically assessed which allows for reliable sensitivity analysis and robustness evaluation taking the variability of the airfoil, profiled endwalls (PEW) at hub and shroud, wedge surfaces (WSF) and the MPCS into account. The probabilistic method — Monte-Carlo simulation (MCS) using an extended Latin Hypercube Sampling (eLHS) technique — is presented subsequently. Afterwards, the FE model that involves thermal, linear-elastic stress and creep analysis is described briefly. Based on this, the fully automated process chain involving CAD model creation, FE mesh morphing, FE analysis and post-processing is executed. Here, the mesh morphing process is presented involving a discussion of the mesh quality. The process robustness is assessed and quantified referring to the impact on input parameter correlation. Finally, the result quantities of the probabilistic FE simulation are evaluated in terms of sensitivities. For this purpose, regions of interest are determined, wherein the statistical analysis is conducted to achieve the sensitivity ranking. A significant influence of the considered geometric uncertainties onto mechanical output quantities is observed which motivates to incorporate these in modern design strategies or robust optimization.

Abstract Synthetic gas, syngas, is a popular alternative fuel for the gas turbine industry. However, the composition of syngas can contain different types and amounts of contaminates, such as carbon dioxide, moisture, and nitrogen, depending on the industrial process involved in its manufacturing. The presence of steam in syngas blends is of particular interest from a thermo-chemical perspective as there is limited information available in the literature. This study investigated the effect of moisture content (0–15% by volume), temperature (323–423 K), and pressure (1–10 atm) on syngas mixtures by measuring the laminar flame speed in a constant-volume, heated experimental facility. A design-of-experiments methodology was applied to these variables to efficiently cover the widest range of conditions that are relevant to the gas turbine industry. The experimental flame speed data were compared to a recent chemical kinetics model showing good agreement. Also, the measured Markstein lengths of atmospheric mixtures were compared and demonstrate that temperature and steam dilution have weak impacts on the sensitivity to stretch in comparison with an increase of carbon monoxide concentration. Mixtures with high levels of CO stabilize the flame structure of thermal-diffusive instability. The increase of steam dilution has only a small effect on the laminar flame speed of high-CO mixtures, while more hydrogen-dominated mixtures demonstrate a much larger and negative effect of increasing water content on the laminar flame speed.

Abstract For gas turbines with free power turbines, the capacity or flow parameter matching is of prime importance. Accurately matched capacity enables the gas turbine to run at its optimum conditions. This ensures maximum component efficiencies, and optimum shaft speeds within mechanical limits. This paper presents the challenges, uncertainties, and opportunities associated with an accurate matching of a generic two-shaft aeroderivative HP-LP gas generator with the free power turbine. Additionally, generic performance trends, uncertainty quantification, and results from the verification program are also discussed. These results are necessary to ensure that the final free power turbine capacity is within the allowable range and hence the product meets the performance guarantees. The sensitivity of free power turbine capacity to various design variables such as the vane throat area, vane trailing edge size, and manufacturing tolerance is presented. In addition, issues that may arise due to not meeting the target capacity are also discussed. To conclude, in addition to design, analysis, and statistical studies, a system-of-systems approach is mandatory to meet the allowed variation in the free power turbine capacity and hence the desired gas turbine performance.

Abstract Smaller combustion chambers, as well as pressure gain combustion processes, are desired in terms of reducing the space, weight, lowering engine manufacturing costs, and increasing the efficiency and specific power of gas turbine engines. In this paper, a compact single rotor disk pressure gain combustor engine is introduced with interesting merits compared with existing turbine engine designs. The engine introduced as the Radial Wave Engine (RWE) features purely radial-flow resulting in a compact and light weight design with only two moving parts. The core of the engine is a spinning radial combustor in which constant volume combustion occurs using stationary inlet and exit end walls resembling a valved-combustor. Power output is generated via a radial-outflow turbine forming a purely disk-shape engine. The proposed engine is a hybrid engine concept, midways between piston and turbine engines with improvements to other counterpart designs. This paper discusses the engine background, operating principles, instrumentation, high-speed video recording, and initial testing of a prototype proof-of-concept demonstrator. This study is the first step towards a new pressure gain combustion configuration which has not designed or tested before.

Abstract The use of metallic Additive Layer Manufacturing (ALM) is an active area of development for gas turbine components, particularly concerning novel combustor prototypes for micro gas turbines. However, further study is required to understand the influence of this manufacturing technique and subsequent post-processing on the resulting burner component surface roughness and its effect on flame stability. In this study, two Inconel 625 swirl nozzle inserts with identical bulk geometry (swirl number, S g = 0.8) were constructed via ALM for use in a generic gas turbine swirl burner. Further post-processing by grit blasting of one swirl nozzle insert results in a quantifiable change to the surface roughness characteristics in the burner exit nozzle when compared with the unprocessed ALM swirl nozzle insert or a third nozzle insert which has been manufactured using traditional machining methods. An evaluation of the influence of variable surface roughness effects from these swirl nozzle inserts is therefore performed under preheated isothermal and combustion conditions for premixed methane-air flames at thermal power of 25 kW. High-speed velocimetry at the swirler exit under isothermal air flow conditions gives evidence of the change in near-wall boundary layer thickness and turbulent fluctuations resulting from the change in nozzle surface roughness. Under atmospheric combustion conditions, this influence is further quantified using a combination of dynamic pressure, high-speed OH* chemiluminescence, and exhaust gas emissions measurements to evaluate the flame stabilization mechanisms at the lean blowoff and rich stability limits. Notable differences in flame stabilization are evident as the surface roughness is varied, and changes in rich stability limit were investigated in relation to changes in the near-wall turbulence intensity. Results show the viability of using ALM swirl nozzles in lean premixed gas turbine combustion. Furthermore, precise control of in-process or post-process surface roughness of wetted surfaces can positively influence burner stability limits and must therefore be carefully considered in the ALM burner design process as well as CFD models.

Abstract Vaporization injectors have been in existence for decades and are a well-proven method of preparing liquid fuels for combustion by heating them above the boiling point of their heaviest hydrocarbon ingredient. By doing so, it converts the fuels into a vapour prior to combustion. When attempting to apply this method of fuel vaporization to micro gas turbines, manufacturing difficulties arise, due to the small complex passages that are required to direct the fuel closer to the high-temperature zone in the combustion chamber and then back to a favourable injection location. This is where the use of additive manufacturing (AM) can prove advantageous due to the complex designs that can be achieved at much smaller scales and potentially at cheaper costs when compared to traditional subtractive manufacturing. The motivation behind the research is to improve the overall efficiency of micro-gas turbines, so they can be applied as range extenders in electric vehicles. Due to the increasing adoption of vehicle electrification. This paper covers the comparison of experimental results for two traditionally manufactured injectors and a third selective laser melted injector (SLM), which were tested in a swirl stabilised micro gas turbine can type combustor on the University of Baths gas stand. The operating range of the tests was 1–4 Bar and 30 to 630 °C inlet air. To the authors knowledge, this is the first such comparison to be made for a gas turbine in open literature, despite wide reports of AM being used in large gas turbines. From the tests, it was found that the 3 and 8 hole machined injectors could not produce stable combustion at the desired operating condition of 4 Bar and 630 °C. The SLM 8 hole injector, however, was able to sustain a stable and constant burn at this design point with low NO x , CO and THC emissions. It was also noticed that the flame colour changed from a yellow flame when testing the first two injectors, to a blue flame when testing the SLM injector suggesting more complete combustion was being achieved due to the lack of soot in burned products, this was assumed to be due to the fuel reaching its saturation conditions within the injector. A number of measurements were taken at various points around the combustor, which included temperatures, pressures and emissions readings. These results were then used to create and validate a non-premixed steady diffusion flamelet model in ANSYS Fluent for the AM injector case. The CFD results were found to overpredict the temperature by approximately 10% when compared to the thermocouple values. This was found to be similar to other studies with similar experimental and computational setups, so it was deemed acceptable. From the validated CFD model, the heat flux at the front surface of the injector was extracted, to be used in a simple heat balance model. Based on a conservative estimate of fuel temperature, the model found that the SLM injectors should have created very near saturation conditions in the nozzle. As this was a conservative analysis, it confirms the experimental findings that partially vaporized fuel was exciting the injector. The model also showed that the fuel in the traditionally machined, 8 hole injector would most likely exit as a liquid.

Abstract Additive manufacturing (AM) is a promising technology for producing better burners. Achieving better energy efficiency on a system level (CO 2 emissions) and lower NOx, particulate emissions and CO, as directed by the International Civil Aviation Organisation (ICAO) standards, is a priority for all aircraft and aircraft engine manufacturers. At the current state-of-the-art, technologies like Powder Bed Fusion (PBF) offer a certain freedom of design one can make good use of. Instead of starting from an established conventional burner design and improve it using AM, the proposed approach in this paper is to define from scratch a design that maximises the potential benefits of AM towards a better burner. However, there are a few playing rules one must be aware of. The design, manufacturing and testing of a staged premixed burner with separate injection ramp was done as the follow-up of paper GT2018-75165 where new swirler shapes had been assessed. For this paper a monolithic, profiled burner design for premixed injection was tested for low-emission combustion. Additional features were included and assessed. Separately, regarding the fuel injection system a new design of a fuel ramp disconnected from the burner is proposed in a first approach, which combines the injection and pre-heating of the fuel. It serves as a fuel splitter (burning fuel / bypassed flow), as a miniature heat exchanger and as a multipoint injection ramp. A merging of the monolithic burner and the injection ramp is planned at a later stage. The fuel injection system using AM parts is assessed separately from the burner in a first approach. It suggests some novel technical solutions regarding 3D printed burner designs. Early combustion experiments are described and supported with function tests using a carefully selected instrumentation.

Abstract With the recent advancement of metallic additive manufacturing (AM), it is perceived that future gas turbines will be manufactured with significantly fewer parts, leading to both financial and safety improvements achieved from reductions in weight, assembly processes and failure modes associated with welded parts. In addition the design and manufacture of highly intricate parts such as fuel atomizers become free from the constraints of tooling, facilitating more complex internal flow geometries to be conceived which afford improved atomization, flame stability and hence combustion efficiency. However, it is noted that increased dimensional tolerances and surface roughness resulting from this manufacturing technique can detrimentally impact internal air and fuel flow paths and hence warrant further investigation. In this study a small-scale (200kW) pre-filming airblast atomizer, based on the Parker Hannifin commercial concept, and typical of injectors utilized in RQL aviation combustors, was manufactured by Cardiff School of Engineering’s High Value Manufacturing Laboratories. Direct metal laser sintering, was utilized to produce a fully operational single component part, manufactured in 316-grade stainless steel using a Renishaw AM250 system, providing a part with measured surface roughness (Ra) values of 12–26 μm in agreement with expected values reported in the literature. Operation of the injector as a single fluid atomizer demonstrated that the fuel channel and integrated swirlers were sufficiently accurate and concentric to result in a uniform spray pattern, displaying global liquid sheet structures which were in agreement with those previously reported. However, the effective area of the atomizer’s air-flow path, when evaluated using differential pressure measurements, was shown to be smaller than predicted, resulting in an increased pressure drop. Laser diffraction droplet sizing was utilized to evaluate the global SMD of the prefilming airblast water spray at atmospheric conditions, across a range of air to liquid ratios. SMD’s between 4.2–115μm were measured at corresponding air-flow rates of 3–25 g/s, with droplet sizes observed to decrease exponentially at higher air-flow rates. This data is again in excellent agreement with SMD correlations previously proposed. Flame stability experiments conducted at ambient pressure and elevated air temperature, demonstrated the stability of a conventional (JET A-1) fuel flame across a range of air and fuel flow rates, representative of pressure drops and AFRs in commercial operation. Further post-processing of the internal flow path walls and swirl vanes to reduce surface roughness is anticipated to result in a lower pressure drop across the air-path geometry, highlighting the potential for further improvements in AM injector performance.