This paper describes a framework to quantify the effect of freestream turbulence, generated by mixing processes in a combustor, on turbulent boundary layer loss generation in the high-pressure turbine downstream of the combustor. The regime of freestream turbulence common to gas turbine aero engines is identified and it is shown that the dissipation loss coefficient in this regime can be determined using existing measurements of the effect of freestream turbulence on skin friction. The paper shows that combustor-generated freestream turbulence can increase the profile loss coefficient of a typical high-pressure turbine blade by as much as 28%. A relation has been derived between a non-dimensional turbulence parameter, which characterizes the freestream turbulence, and the increase in turbine boundary layer dissipation, which quantifies the decrease in turbine efficiency. The relation provides guidelines for combustor turbulence modifications that lead to turbine performance benefits. The framework has been applied in example trade studies which show that increasing the size of dilution ports and increasing the length of the combustor can decrease high-pressure turbine profile loss generation to potentially increase stage efficiency up to 0.5%.

This paper describes the conceptual design of a boundary layer ingesting, distributed electric propulsor system for civil aircraft, supported through NASA's CHEETA (Center for High-Efficiency Electrical Technologies for Aircraft) program. Cryogenic hydrogen fuel cell-powered superconducting electric motors power the boundary layer ingesting ducted fans to achieve high propulsive efficiency. Sizing of the integrated propulsor system is carried out using a Geometric Programming (GP) optimization framework to capture the effects of system integration on component design. A 16-propulsor configuration is presented which can meet the requirements of the CHEETA aircraft using motors similar to a previously reported 2.5-MW reference design. Analysis of configurations with different numbers of propulsors shows that the aero-propulsive benefit of distributed propulsion is limited by the thermal management requirements and constraints on the geometry of the superconducting machine which do not scale linearly with shaft power. The resulting fan and motor designs are different than if each were designed independently, illustrating the importance of simultaneous optimization in the design of novel electric propulsors.

This paper describes the role of tip leakage flow in creating the leading edge separation necessary for the onset of spike-type compressor rotating stall. A series of unsteady multipassage simulations, supported by experimental data, are used to define and illustrate the two competing mechanisms that cause the high incidence responsible for this separation: blockage from a casing-suction-surface corner separation and forward spillage of the tip leakage jet. The axial momentum flux in the tip leakage flow determines which mechanism dominates. At zero tip clearance, corner separation blockage dominates. As clearance is increased, the leakage flow reduces blockage, moving the stall flow coefficient to lower flow, i.e., giving a larger unstalled flow range. Increased clearance, however, means increased leakage jet momentum and contribution to leakage jet spillage. There is thus a clearance above which jet spillage dominates in creating incidence, so the stall flow coefficient increases and flow range decreases with clearance. As a consequence, there is a clearance for maximum flow range; for the two rotors in this study, the value was approximately 0.5% chord. The chordwise distribution of the leakage axial momentum is also important in determining stall onset. Shifting the distribution toward the trailing edge increases flow range for a leakage jet dominated geometry and reduces flow range for a corner separation dominated geometry. Guidelines are developed for flow range enhancement through control of tip leakage flow axial momentum magnitude and distribution. An example is given of how this might be achieved.

William Rede Hawthorne was a pioneer in gas turbine aerodynamics and thermodynamics, a sought-after technology advisor to industry and government, and a generous and enthusiastic teacher who encouraged students to excel. His outstanding contributions included resolution of combustion problems that limited the operation of the original Whittle jet engine, early in-depth descriptions of compressible channel flow that still inform engineers today, innovative and wide-ranging analyses of secondary flows in turbomachinery that defined the field, and creation of some of the first notes on gas turbine cycle analysis. A theme in the many areas of engineering in which he had impact was the satisfaction from the growth of understanding that can accompany making things work—in his words, ‘machines produce ideas just as surely as ideas produce machines’. A Cambridge graduate, he was a professor at MIT when, in 1951, he was recruited to a newly established chair at Cambridge, where he later had leadership roles as head of the engineering department (1968–73) and Master of Churchill College (1968–83). He retained strong ties to MIT, however, and fostered lasting collaborations between the two universities. Among his numerous awards and honours were the US Medal of Freedom (1947), a Royal Society Medal (1982) and a knighthood (1970) for ‘services to thermodynamics’, a citation that pleased him greatly.

Propulsors with boundary layer ingestion (BLI) generate a propulsive force with lower flow power input than conventional engines. This aerodynamic benefit can be traced back to its sources: reductions in jet, surface, and wake dissipation. A framework for BLI analysis is developed based on the power balance method: parametric expressions are derived for the net streamwise force on an aircraft and for the mechanical flow power required, as well as relations for the dissipation components. They illustrate the range of possible comparisons between BLI and non-BLI engine installations, and show how the benefit varies with design choices. Applying the framework to wind tunnel data from powered D8 aircraft model tests, the sources of aerodynamic BLI benefit are quantified for a realistic transport aircraft configuration. With the same propulsors (equal nozzle area) BLI reduces cruise power by 8.2%, of which 5.2% comes from a reduction in jet dissipation, 2.4% from reduced surface dissipation, and 0.6% from reduced wake dissipation. The jet dissipation reduction is equivalent to a 3.4 percentage points increase in propulsive efficiency. If the installations are compared at equal mass flow, the benefit amounts to a 9% lower cruise power.

This paper describes a new conceptual framework for three-dimensional turbomachinery flow analysis and its use to assess fan stage attributes for mitigating adverse effects of inlet distortion due to boundary layer ingestion (BLI). A non-axisymmetric throughflow method has been developed to describe the fan flow field with inlet distortion. In this the turbomachinery is modeled using momentum and energy source distributions that are determined as a function of local flow conditions and a specified blade camber surface geometry. Comparison with higher-fidelity computational and experimental results shows that the method captures the principal flow redistribution and distortion transfer effects associated with BLI. Distortion response is assessed for a range of (i) rotor spanwise work profiles, (ii) rotor-stator spacings, and (iii) non-axisymmetric stator geometries. For the parameters examined, changes in axisymmetric design result in trades between rotor and stator distortions, or between different radial sections of a given blade row with marginal overall gain. Of the approaches examined, non-axisymmetric stator exit flow angle distributions were found to provide the greatest reduction in rotor flow distortion and thus may offer the most potential for mitigating decreases in performance due to BLI inlet distortion.

This paper presents a description of the physical principles of aerodynamic power savings from boundary layer ingestion propulsion and a quantitative evaluation of the boundary layer ingestion benefit for advanced civil aircraft. Control volume and one-dimensional analyses are used to illustrate two major features of boundary layer ingestion: reduction of jet mixing losses due to decreased jet kinetic energy from reduced velocity of flow entering the propulsor and, to a lesser extent, reduction of airframe wake mixing losses. Embedded boundary layer ingestion propulsion systems can also enable nacelles with reduced surface area and associated weight and drag, further decreasing the aircraft propulsive power requirement. The required propulsor flow power is shown to decrease with increases in both the amount of boundary layer ingested and the propulsor mass flow, and there is thus no unique way to compare boundary layer ingestion and non-boundary-layer-ingestion systems. Using the ideas presented, however, the benefit can be assessed for any given comparison. The analysis is applied to an advanced civil transport aircraft concept with 40% of the fuselage boundary layer ingested, yielding a reduction in required propulsor mechanical power of 9% relative to a non-boundary-layer-ingestion configuration with the same propulsors, in agreement with computational fluid dynamics calculations and wind tunnel experiments.

In this paper, we describe the structures that produce a spike-type route to rotating stall and explain the physical mechanism for their formation. The descriptions and explanations are based on numerical simulations, complemented and corroborated by experiments. It is found that spikes are caused by a separation at the leading edge due to high incidence. The separation gives rise to shedding of vorticity from the leading edge and the consequent formation of vortices that span between the suction surface and the casing. As seen in the rotor frame of reference, near the casing the vortex convects toward the pressure surface of the adjacent blade. The approach of the vortex to the adjacent blade triggers a separation on that blade so the structure propagates. The above sequence of events constitutes a spike. The computed structure of the spike is shown to be consistent with rotor leading edge pressure measurements from the casing of several compressors: the centre of the vortex is responsible for a pressure drop and the partially blocked passages associated with leading edge separations produce a pressure rise. The simulations show leading edge separation and shed vortices over a range of tip clearances including zero. The implication, in accord with recent experimental findings, is that they are not part of the tip clearance vortex. Although the computations always show high incidence to be the cause of the spike, the conditions that give rise to this incidence (e.g., blockage from a corner separation or the tip leakage jet from the adjacent blade) do depend on the details of the compressor.

This paper describes a rotating rake system for rapid measurement of total and static pressure distributions in powered airframe models with complex geometries. The local actuation of the rake rotation mechanism allows access to regions of the flow not easily accessible to traditional traverse systems with wall-mounted leadscrew actuators. Local indexing of the rake pivot to the model precisely positions the rake in the presence of model vibration. The total and static rakes and mechanisms are built from commercially available hardware and materials. The rotating rake system was employed to measure total and static pressures in the propulsive streamtubes of a boundary-layer ingesting powered aircraft model in the NASA Langley 14’×22’ wind tunnel. The specific quantity of interest was the total pressure flux into and out of the propulsors, needed to quantify the power-saving benefits of the boundary layer ingestion system. The paper discusses the benefits and drawbacks of the rotating rake system for the demonstrated wind tunnel experiment and for more general flow measurements.

This paper presents a framework for estimating the upper limit of compressor stage efficiency. Using a compressor stage model with a representative design velocity distribution with turbulent boundary layers, losses are calculated as the sum of selected local irreversibilities, rather than from correlations based on data from existing machines. By considering only losses that cannot be eliminated and optimizing stage design variables for minimum loss, an upper bound on stage efficiency can be determined as a function of a small number of stage design parameters. The impact of the stage analysis results are evaluated in the context of gas turbine cycle performance. The implication from the results of the stage level and cycle analyses is that compressor efficiency improvements that result in substantial increases in cycle thermal efficiency are still to be realized.

The effect on rotor work, of the phase of an upstream wake relative to the rotor is examined computationally and analytically for a transonic blade row. There can be an important impact on time-mean performance when the time-dependent circulation of the shed vortices in the wake is phase-locked to the rotor position, as occurs when there is strong interaction between rotor static pressure field and upstream vanes. The rotor work is found to depend on the path of the wake vortices as they travel through the blade passage; for configurations examined, the calculated change in time-mean rotor work was approximately three percent. It is shown that the effect on work input can be analyzed in terms of the influence of the time-mean relative stagnation pressure nonuniformity associated with the unsteady (but phase-locked) wake vortex flow field, in that changes in vortex path alter the location of the nonuniformity relative to the rotor Lower pressure rise and work input occurs when the rotor blade is embedded in a region of low time-mean relative stagnation pressure than when immersed in a region of high relative stagnation pressure. In addition to the work changes, which are an essentially two-dimensional effect, it is demonstrated that the location of the wake may affect the tip clearance flow, implying a potential impact on pressure rise capability and rotor stability limits. Model calculations are presented to give estimates of the magnitude and nature of this phenomenon.

This paper presents an assessment of the performance of an embedded propulsion system in the presence of distortion associated with boundary layer ingestion. For fan pressure ratios of interest for civil transports, the benefits of boundary layer ingestion are shown to be very sensitive to the magnitude of fan and duct losses. The distortion transfer across the fan, basically the comparison of the stagnation pressure non-uniformity downstream of the fan to that upstream of the fan, has a major role in determining the impact of boundary layer ingestion on overall fuel burn. This, in turn, puts requirements on the fidelity with which one needs to assess the distortion transfer, and thus the type of models that need to be used in such assessment. For the three-dimensional distortions associated with fuselage boundary layers ingested into a subsonic diffusing inlet, it is found that boundary layer ingestion can provide decreases in fuel burn of several per cent. It is also shown that a promising avenue for mitigating the risks (aerodynamic as well as aeromechanical) in boundary layer ingestion is to mix out the flow before it reaches the engine face.

Problems of high technological interest, for example the development of gas turbine engines, span disciplinary boundaries. Collaboration is critical in advancing the technology, but it has been less a factor in gas turbine research. In this paper it is proposed that step changes in gas turbine performance can emerge from such collaborative endeavors. In these, success depends on the development of integrated teams with the appropriate range of skills. This is well known in product development, but it is less familiar, and less subscribed to, in the research community. Case histories of two projects are given to illustrate the point: the development of the concept of “smart jet engines” and the Silent Aircraft Initiative. In addition to providing the ability to attack multidisciplinary issues discussion is also given about the way in which collaboration can enhance the research process within a single discipline.

This paper presents a new criterion for estimating the onset of three-dimensional hub-corner stall in axial compressor rotors and shrouded stators. A simple first-of-a-kind description of hub-corner stall formation is developed which consists of (i) a stall indicator, which quantifies the extent of the separated region via the local blade loading and thus indicates whether hub-corner stall occurs, and (ii) a diffusion parameter, which defines the diffusion limit for unstalled operation. The stall indicator can be cast in terms of a Zweifel loading coefficient. The diffusion parameter is based on preliminary design flow variables and geometry. Computational simulations and single and multistage compressor data are used to show the applicability of the criterion over a range of blade design parameters. The criterion also enables determination of specific flow control actions to mitigate hub-corner stall. As an illustration, a flow control blade, designed using the ideas developed, is seen to produce a substantial reduction in the flow nonuniformity associated with hub-corner stall.

Subtitle: Working with partners in industry, university research departments hope to conceive an aircraft design whose engine and airframe noise would be imperceptible in the urban environment around airports.

This Brief Communication describes features of steady irrotational flow in a viscous fluid. The aim is to illustrate, in a simple manner, the different roles played by viscous forces (which are identically zero) and viscous stresses (which are not), the differences in behavior of stagnation pressure (which is constant along a streamline even in the presence of viscous dissipation) and stagnation enthalpy (which changes through the flow), and the role of viscous stresses on the boundary in creating these behaviors.

This report discusses the results of a study mandated by the U.S. Congress in 2003 to examine and mitigate the environmental impacts of aircraft noise and exhaust gases; and to increase the fuel efficiency of aircraft. The report is divided into six sections, which present the relationship between the environment and aviation, and provide recommendations and a national vision.

Most accounts of the history of turbomachinery look at the machines that were produced. This paper looks at the underlying ideas behind the designs and the methods that were used. It will be seen that in this history the inventiveness of engine designers far outstripped the capabilities of these methods and, further, that this gap spurred the advancement of new ideas. Thus, although analyses were unable to capture many features of the flows, this has not precluded aeroengine turbomachinery from being successfully developed using methods that were far from a complete description. The paper concludes with a look at two different ways in which engineers have dealt with some turbomachinery aerodynamic issues that seemed to offer major difficulties at the time.