Experimental and Numerical Performance Analysis of a Self Starting, Three-Dimensional SCRamjet Intake

In the present paper, a three-dimensional intake model is investigated in a blow-down wind tunnel and results are compared and complemented with numerical simulations (Reynolds-averaged Navier-Stokes simulations). The intake model was equipped with a movable cowl and therefore the internal contraction ratio was variable and was adjusted to the self starting condition. Three different conditions were investigated: i) at a free stream Mach number of 7, a v-shaped cowl geometry was investigated, ii) at a free stream Mach number of 7 a straight cowl geometry was investigated, and iii) the v-shaped cowl geometry was investigated at a free stream Mach number of 6. Furthermore a one-dimensional post analysis was performed, to calculate overall engine parameter from stream thrust averaged intake performance parameter. The numerical simulations were validated on the basis of wall pressure and rake pressure measurements and the match was generally well. Performance wise, the v-shaped cowl was slightly superior compared to the straight cowl. For the Mach 7 configurations maximum back pressure ratios of approximately 120 were measured, which is about 4 times the nominal operating pressure ratio. For the Mach 6 case, the maximum static back pressure ratio dropped to about 80, which again is approximately 4 times the nominal operating pressure ratio

Modified Kantrowitz Starting Criteria for Mixed Compression Supersonic Intakes

A proper understanding of the intake starting process of supersonic airbreathing engines is crucial for a successful operation of the supersonic aircraft, especially for three-dimensional geometries in which no reliable starting prediction for a broad Mach number range exists, and the widespread Kantrowitz criterion only provides a conservative prediction. Experimental investigations from the literature are reviewed and put into the perspective of the Kantrowitz theory. First, an empirical relation is developed that is valid for a wide Mach number range, and that can be calibrated to certain classes of intakes by the user. Second, the Kantrowitz assumptions are modified and a semiempirical relation is derived. The semiempirical relation turned out to be an optimistic limit for self-starting

Experimental Investigation of the Starting Behavior of a Three-Dimensional Scramjet Intake

To this date, there is no generally valid method to accurately predict intake starting for a certain three dimensional scramjet intake configuration. Therefore, experiments are conducted at Mach6 and 7 in a blowdown wind tunnel on a three-dimensional intake model, equipped with a movable cowl and therefore variable internal contraction. The internal contraction is slowly decreased, and the performance parameters at the interface to the combustion chamber are measured with a rake. The general trends from the Kantrowitz diagram were able to be reproduced, namely enhanced intake starting for higher freestream Mach numbers and lower internal contraction ratios. Furthermore, the effect of the angle of attack was twofold: When the intake was pitched, the leading-edge shock decreased in strength, and the intake showed improved starting characteristics. For increasing yaw angles, intake starting was hindered. Exchanging the V-shaped cowl with a straight-cowl geometry improved intake starting due to the additional mass spillage. Once the intake was started, increasing the internal contraction ratio to 2.56 did not cause the intake to unstart for all Mach numbers investigated

Viscous Effects and Truncation Effects in Axisymmetric Busemann SCRamjet Intakes

While designing a Busemann intake, we encountered two main constraints: First, caused by the displacement of the boundary layer, the pressure levels in a Busemann intake were higher (40%) than the values predicted analytically; these effects were labeled viscous effects. Second, truncating the flow field also changed the properties of the classical Busemann flow; these effects were labeled truncation effects and we distinguished between truncating the intake at the leading edge and at the rear side. To take into account viscous effects, we calculated the boundary layer displacement thickness with an integral method and widened the inviscid Busemann contour. To quantify the leading edge truncation effect, an oblique shock at the leading edge was assumed, and also the intake contraction ratio was stretched. The change in flow properties during the rear side truncation was derived in a stream thrust analysis. We separately investigated the effects with Reynoldsaveraged Navier-Stokes and Euler simulations, respectively. After the viscous correction, the deviation between numerically and analytically calculated pressure levels was below 10%. When truncating at the leading edge, the numerically calculated pressure level was generally too low, while the temperature level was within 5% of the prediction. After stretching the contraction ratio, we were able to keep the pressure within 5% of the design value, however in this case, the temperatures were generally too high. The stream thrust approach was highly accurate and for rear side truncated intakes we were able to predict the numerical pressure and temperature levels within 1% of the design values

Design and Performance Analysis of Three-Dimensional Air Intakes for Supersonic Combustion Ramjet Engines

The Supersonic Combustion Ramjet (SCRamjet) engine is an efficient propulsion device for supersonic and hypersonic velocities. The present work focuses on the SCRamjet air intake which serves as the engine’s compressor. First, an analytical intake design tool was developed to generate so called streamline traced intake geometries. Second, the starting behavior of hypersonic air intakes was investigated experimentally in a blow down wind tunnel with a three-dimensional intake model with planar surfaces. A semi-empirical estimate was developed to predict intake starting of three-dimensional intake geometries. Third, the three-dimensional intake was further investigated experimentally and results were compared to numerical simulations. Finally, the performance of the three-dimensional intake was compared to the performance of different streamline traced intakes, designed with the analytical tool. Overall, the specific impulse of the streamline traced intakes was approximately 10 % higher compared to the three-dimensional intake configuration

Axisymmetric SCRamjet Engine Design and Performance Analysis

In the present work, we treat the SCRamjet engine as a system, by coupling different analytical design tools. For the intake portion a Busemann tool was used, the combustion chamber was modeled with a one dimensional reactor, and the nozzle was modeled with the method of characteristics. Performance parameter are discussed while we mainly focus on specific impulse. We used numerical simulations to validate the modeling of the combustion chamber and it turned out that after analytically forcing ignition, the numerical results matched the predictions best. Furthermore, we validated the analytical approach with numerical results and obtained good agreement. According to the analytical tool, for a constant intake pressure ratio of 80 the engine did not ignite at Mach numbers lower than 6 and the specific impulse dropped from 3500 s at Mach 6 to 1200 s and 500 s for Mach 8 and 10, respectively. Due to the small geometry and therefore high friction drag, specific impulse became negative at Mach 11

Converging-Diverging Nozzles with Constant-Radius Centerbody

Several flow phenomena, such as recirculating wake flows or noise generation, occur in aerodynamic configurations with backward facing steps. In this context subsonic nozzles with constant-radius centerbodies exist, which enable fundamental research of these phenomena for M < 1. For the supersonic regime, however, the existing database and knowledge is limited. Therefore, the present work presents a design approach for a converging-diverging nozzle with constant-radius centerbody. For the nozzle throat, Sauer’s method is modified to include a centerbody. The method of characteristics is used for the subsequent supersonic portion. Comparing the analytical calculations to numerical simulations results in very good agreement and therefore underlines the feasibility of the chosen approach. Viscosity reduced the Mach number on the exit plane by 1.0 − 1.2% and therefore had little influence

Experimental and Numerical Performance Analysis of a Self-Starting Three-Dimensional Scramjet Intake

In this work, a three-dimensional intake model was investigated in a blowdown wind tunnel and results were compared and complemented with Reynolds-averaged Navier–Stokes simulations. The intake model was equipped with a movable cowl with which the internal contraction ratio was set to the self-starting limit. Three different conditions were investigated: first, a v-shaped cowl geometry at a freestream Mach number of seven; second, a straight cowl geometry at a freestream Mach number of seven; and, finally, a v-shaped cowl geometry at a freestream Mach number of six. Furthermore, a one-dimensional postanalysis was performed to calculate overall engine parameters from stream-thrust-averaged intake performance parameters. Numerical results were within the experimental uncertainty, except for small displacements near separation regions. The mass capture ratio of the v-shaped cowl was slightly higher as compared to the straight cowl. When increasing the Mach number from six to seven, the specific impulse dropped from 2561 to 2100 s, respectively. For the Mach 7 configurations, a Maximum sustainable backpressure ratio of approximately 115 was measured. For the Mach 6 case, the maximum sustainable static backpressure ratio dropped to about 85. The maximum sustainable backpressure to operating backpressure ratio was around 4:1, and it was independent of the Mach number

Design of converging-diverging nozzles with constant-radius centerbody

Several flow phenomena, such as recirculating wake flows or noise generation, occur in aerodynamic configurations with backward facing steps. In this context, subsonic nozzles with constant-radius centerbodies exist, which enable fundamental research of these phenomena for M < 1. For the supersonic regime, however, the existing database and knowledge are limited. Therefore, this work presents a design approach for a converging-diverging nozzle with constant-radius centerbody. For the nozzle throat, Sauer’s method is modified to include a centerbody. The method of characteristics is used for the subsequent supersonic portion. Comparing the analytical calculations to numerical simulations results in very good agreement and therefore underlines the feasibility of the chosen approach. Viscosity reduced the Mach number on the exit plane by 1.0–1.2% and therefore had little influence