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A topic of increasing importance in the Unmanned Aerial System (UAS) community is the design and performance of open propellers used in hand launched, small UASs. The design and testing of these propellers is necessary to accurately predict UAS operation. This paper describes the design methodology used by Baylor University and the USAF Academy to design propeller blades for optimum efficiency and low noise. Propeller blade design theories are discussed as well as an overview of several of the existing design codes. Included is a discussion on geometric angle of attack, the induced angle of attack, and their impact on propeller design. The design program BEARCONTROL was developed which incorporates the programs QMIL and QPROI). Supplemental codes were also developed to work with Bearcontrol to design a propeller with a constant chord and variable twist. This resulted in the angle of attack for L/Dmax being used from the propeller hub to the tip. BEARCONTROL is a program written in MATLAB that gives a user the ability to quickly design a propeller, predict its performance, and then create a 30 model in SolidWorks. The MATLAB GUl ultimately results in a mostly automated process that is simple to use for individuals who are unfamiliar with command prompt programs and SolidWorks modeling. Also incorporated into BEARCONTROL is the program NREL AirFoil Noise (NAFNOISE) developed by the National Renewable Energy Laboratory (NREL). This program predicts the noise of any airfoil shape and provides a comparison for optimizing/minimizing predicted noise for the propeller being designed. Construction methods and materials also have a direct impact on cost, durability and operability when using a rapid prototype process to fabricate propellers. An overview of materials and construction methods used in this research are discussed. Incorporation of a hub with interchangeable blades is also presented as a more efficient testing method.

A topic of increasing importance in the Unmanned Aerial System (UAS) community is the design and performance of open propellers used in hand launched, small UASs. The performance of these small propellers directly influences the operational capabilities of the UAS. As such, the design and testing of these propellers is necessary to accurately predict UAS performance. This experimental investigation examined the relationship between diameter, pitch, and number of blades to aerodynamic efficiency and aero-acoustic sound pressure levels. Thrust, torque, propeller rotational speed, and sound pressure level were measured for twelve aero-nautCAMcarbon (ACC) folding propeller configurations currently being used on an operational UAS with diameters ranging from 12 to 15 inches, pitches from 6 to 13 inches and increasing from two to three propeller blades. Each configuration was tested at 44 ft/s tunnel velocity, the typical cruising velocity of a small UAS, while the propeller rotational speed was varied to determine the rotational speed needed to produce 2.5 lbf of thrust, a typical cruise thrust required for a small UAS. As expected, the rotational speed required to achieve the desired thrust decreased approximately 7.7% per inch increase in the propeller diameter. At the same time, the noise signature decreased by approximately 0.8 dB per inch increase and overall efficiency rose by 2.9% per inch increase. Similar results were found for increasing both the number of propeller blades and also increasing the pitch of the propeller. Increasing from two to three blades decreased the rotational speed by 9.1% with a 2.1 dB drop in sound pressure level and an increase in overall efficiency of 3.2%. Increasing the pitch generally decreased rotational speed by 4.6% per inch of pitch increase and decreased noise level by 0.7dB per inch of pitch increase. Overall efficiency slightly increased by 0.6% for an inch increase in pitch. For a given diameter propeller there seems to be an optimum pitch for minimum sound pressure level. For design, this indicates there is an optimum angle of attack for the propeller, which translates to an optimum beta twist angle to achieve minimum sound pressure level. Noise generation was found to be a strong function of propeller rotational speed. Lower rotational speed generally produced less noise.

Pulse detonation engines (PDEs) are currently attracting considerable research and development attention because they promise performance improvements over existing airbreathing propulsion devices. Because of their inherently unsteady behavior, it has been difficult to conveniently classify and evaluate them relative to their steady-state counterparts. Consequently, most PDE studies employ unsteady gasdynamic calculations to determine the instantaneous pressures and forces acting on the surfaces of the device and integrate them over a cycle to determine thrust performance. A classical, closed thermodynamic cycle analysis of the PDE that is independent of time is presented. The most important result is the thermal efficiency of the PDE cycle, or the fraction of the heating value of the fuel that is converted to work that can be used to produce thrust. The cycle thermal efficiency is then used to find all of the traditional propulsion performance measures. The benefits of this approach are 1) that the fundamental processes incorporated in PDEs are clarified; 2) that direct, quantitative comparisons with other cycles (e.g., Brayton or Humphrey) are easily made; 3) that the influence of the entire ranges of the main parameters that influence PDE performance are easily explored; 4) that the ideal or upper limit of PDE performance capability is quantitatively established; and 5) that this analysis provides a basic building block for more complex PDE cycles. A comparison of cycle performance is made for ideal and real PDE, Brayton, and Humphrey cycles, utilizing realistic component loss models. The results show that the real PDE cycle has better performance than the real Brayton cycle only for flight Mach numbers less than about 3, or cycle static temperature ratios less than about 3. For flight Mach numbers greater than 3, the real Brayton cycle has better performance, and the real Humphrey cycle is an overoptimistic (and unnecessary) surrogate for the real PDE cycle.

mi = initial mass of the space transportation system or stage, lbm mp = orbital payload mass of the space transportation system, lbm Z = ratio of the initial mass of an airbreathing single-stage-to-orbit system to that of an airbreathing two-stage-to-orbit system of equal technological capability Ze = ratio of the empty mass of an airbreathing single-stage-to-orbit system to that of an airbreathing two-stage-to-orbit system of equal technological capability = ratio of initial space transportation system mass to orbital payload mass, or initial mass ratio e = ratio of empty mass to initial mass for a stage f = ratio of fuel plus oxidizer mass to initial mass for a stage

The liquid oil lubrication system of current aircraft jet engines accounts for approximately 10–15% of the total weight of the engine. It has long been a goal of the aircraft gas turbine industry to reduce this weight. Vapor-Phase Lubrication (VPL) is a promising technology to eliminate liquid oil lubrication. The current investigation resulted in the first gas turbine to operate in the absence of conventional liquid lubrication. A phosphate ester, commercially known as DURAD 620B, was chosen for the test. Extensive research at Wright Laboratory demonstrated that this lubricant could reliably lubricate railing element bearings in the gas turbine engine environment. The Allison T63 engine was selected as the test vehicle because of its small size and bearing configuration. Specifically, VPL was evaluated in the number eight bearing because it is located in a relatively hot environment, in line with the combustor discharge, and it can be isolated from the other bearings and the liquid lubrication system. The bearing was fully instrumented and its performance with standard oil lubrication was documented. Results of this baseline study were used to develop a thermodynamic model to predict the bearing temperature with VPL. The engine was then operated at a ground idle condition with VPL with the lubricant misted into the #8 bearing at 13 ml/hr. The bearing temperature stabilized at 283°C within 10 minutes. Engine operation was continued successfully for a total of one hour. No abnormal wear of the rolling contact surfaces was found when the bearing was later examined. Bearing temperatures after engine shutdown indicated the bearing had reached thermodynamic equilibrium with its surroundings during the test. After shutdown bearing temperatures steadily decreased without the soakback effect seen after shutdown in standard lubricated bearings. In contrast, the oil lubricated bearing ran at a considerably lower operating temperature (83°C) and was significantly heated by its surroundings after engine shutdown. In the baseline tests, the final bearing temperatures never reached that of the operating VPL system.Copyright © 1997 by ASME

The application of an ejector fitted with a diffuser is used to transfer momentum from a high speed primary flow to an entrained secondary flow to increase the overall thrust of the system. The increase in thrust is due to a greater propulsive efficiency achieved by trading exit velocity for overall mass flow. The objective of this project was to understand the effects of varying the primary nozzle and ejector exit flow areas on overall ejector/diffuser performance. Three primary flow nozzles were tested in combination with four diffuser exit diameters. Thrust augmentation was found to increase as the ratio of the secondary to primary flow areas increased and as the ratio of diffuser exit area to mixer areas increased. A maximum thrust augmentation of 1.5 was achieved with the smallest diameter primary flow nozzle and the largest diameter diffuser exit.

This paper presents the results of a fundamental, comprehensive, and rigorous analytical and computational examination of the performance of the Brayton propulsion and power cycle employing real air as the working fluid. This approach capitalizes on the benefits inherent in closed cycle thermodynamic reasoning and the behavior of the thermally perfect gas to facilitate analysis. The analysis uses a high fidelity correlation to represent the specific heat at constant pressure of air as a function of temperature and the polytropic efficiency to evaluate the overall efficiency of the adiabatic compression and expansion processes. The analytical results are algebraic, transparent, and easily manipulated, and the computational results present a useful guidance for designers and users. The operating range of design parameters considered covers any current and foreseeable application. The results include some important comparisons with more simplified conventional analyses.

A constant-area diffuser, or 'isolator', is required in both the ramjet and scramjet operating regimes of a dual-mode engine configuration in order to prevent unstarts due to pressure feedback from the combustor. Because the nature of the combustor-isolator interaction is different in the two operational modes, however, attention is presently given to the use of thermal vs kinetic energy coordinates for these interaction processes' visualization. The results of the analysis thus conducted indicate that the isolator requires severe flow separation at combustor entry, and that its entropy-generating characteristics are more severe than an equivalent oblique shock. A constant-area diffuser is only marginally able to contain the equivalent normal shock required for subsonic combustor entry.

Aircraft gas turbine analysis and design pedagogy can be enriched through the incorporation improved preliminary engine thrust and fuel consumption models, novel computer programs for both aircraft system analysis and turbomechanical design, and a new perspective for engine-cycle analysis. Four computer programs have been developed for preliminary engine design; two of these automate aircraft system analysis, while another designs multistage axial-flow compressors and the last designs multistage axial-flow turbines. Student confusion with 'design-point' and 'off-design' concepts is by these means reduced. 6 refs.

T HE authors of this technical note assert, without presenting supporting evidence, that the methods of Mattingly et al. "gave results having more than 10-15% error in fuel flow and thrust, even at cruise flight conditions." These methods have, in fact, been shown to reproduce thrust within a few percent, and fuel consumption within about 5% over a wide range of operation, even for complex dual-spool, mixed

We denote the right hand side of Eq. (1) by J(x), and this can be obtained from flow quantities available from our experiment and is shown on Fig. 5. However, Q" (0, x) will not be calculated from our test results in view of the numerical errors inherent in the double differentiation of data. Estimates of Q"(0, x) can be made by assuming a theoretical model, and its value will depend on the model chosen. By assuming Kleinstein's expression for the fully developed region, one obtains

The design of the cascade hardware and baseline airfoils was previously reported. The test results of baseline airfoil performance in the annular cascade and the status of the other tasks are presented in this report.

A comprehensive three-year program is in progress to investigate methods of improving the performance of fan-drive turbines. The goals of the program are to develop turbine aerodynamic techniques and design procedures for efficient, high work, low pressure turbines by means of analytical studies and cascade testing, and to demonstrate the effectiveness of the techniques by designing and testing a two-stage turbine that meets or exceeds the contract stage work and efficiency goals. The first phase effort defining the preliminary turbine design has been completed and the results have been reported. The second phase consists of a preliminary experimental evaluation which includes establishment of two-dimensional loss levels and three- dimensional flow behavior for both the baseline airfoils and airfoils utilizing boundary layer control devices. The design of the rig hardware and the baseline airfoils which will be evaluated during this second phase are presented in this report. (Author)

A comprehensive, four-phase, three-year program is in progress to investigate methods of improving the performance of fan-drive turbines. The goals of this program are to develop turbine design procedures and aerodynamic techniques for high work, efficient, low-pressure turbines. The first phase effort of defining the preliminary turbine design has been completed and the results were reported. The second phase consists of an experimental evaluation which includes establishment of both two-dimensional loss levels and three-dimensional flow behavior for the baseline airfoils and for airfoil utilizing various boundary layer control methods. The design of the baseline cascade packs and the three-dimensional performance of the baseline airfoils were reported. The test results of the baseline airfoil boundary layer control methods and the performance of the decreased solidity annular cascades are presented in this report.

The effects of a uniform magnetic field normal to the shock front upon shock states and expansion wave properties are theoretically and experimentally investigated for large magnetic-drive fields. One-dimensional theory arid several simplifying assumptions are employed to predict the Important effects. and a simple method of visualizing the possible shock states is obtained. The restrictions upon shock-wave formation imposed by various types of instabilities are considered. A series of experiments has been performed in a magnetically driven shock tube of annular cross section under conditions approaching the assumptions. Some difficulty was encountered in producing well- formed shocks. and a tentative link with the drive-current breakdown characteristics is established. Observations of well-formed shocks indicate close agreement with shock-transition theory. even when the axial magnetic field exceeds the switch-on value. although the expansion wave properties exhibit large departures from the theory. Some estimates of flow adjustments caused by radius effects and leakage through the boundary layer have been made, and are found to contribute to an understanding of the experimental results.

A three-year program was initiated to provide a direct attack on the problem of attaining high efficiency in highly loaded turbine stages. The goals of this program are to develop turbine aerodynamic techniques and design procedures for efficient, high work, low pressure turbines by means of analytical studies and cascade testing, and to demonstrate the effectiveness of the techniques and procedures by designing and testing a two stage turbine that meets or exceeds the contract stage work and efficiency goals. The first phase effort described in this report was directed toward defining a turbine design with the highest inherent resistance to boundary layer separation and to select boundary layer control techniques that are best suited for extending the loading limits of the basic turbine design. (Author)

This paper gives the theory of the motion of a finitely conducting, viscous liquid between long, concentric, rotating cylinders under a radial magnetic field in the case of high Hartmann number. Because the current and vorticity contents of a Hartmann layer are related, it is easy to predict the behaviour in both steady and unsteady states. Experiments with mercury which confirm the theory closely are described. As these were developed for an educational film on magnetohydrodynamics, they have the attraction that the results are immediately obvious to the observer.