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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

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.

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

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)