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All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Maximum orbit plane change with heat-transfer-rate considerations

Two aerodynamic maneuvers are considered for maximizing the plane change of a circular orbit: gliding flight with a maximum thrust segment to regain lost energy (aeroglide) and constant altitude cruise with the thrust being used to cancel the drag and maintain a high energy level (aerocruise). In both cases, the stagnation heating rate is limited. For aeroglide, the controls are the angle of attack, the bank angle, the time at which the burn begins, and the length of the burn. For aerocruise, the maneuver is divided into three segments: descent, cruise, and ascent. During descent the thrust is zero, and the controls are the angle of attack and the bank angle. During cruise, the only control is the assumed-constant angle of attack. During ascent, a maximum thrust segment is used to restore lost energy, and the controls are the angle of attack and bank angle. The optimization problems are solved with a nonlinear programming code known as GRG2. Numerical results for the Maneuverable Re-entry Research Vehicle with a heating-rate limit of 100 Btu/ft(2)-s show that aerocruise gives a maximum plane change of 2 deg, which is only 1 deg larger than that of aeroglide. On the other hand, even though aerocruise requires two thrust levels, the cruise characteristics of constant altitude, velocity, thrust, and angle of attack are easy to control.

Lee, J. Y.↗
NASA NTRS2014

Performance of a Supersonic Over-Wing Inlet with Application to a Low-Sonic-Boom Aircraft

Development of commercial supersonic aircraft has been hindered by many related factors including fuel-efficiency, economics, and sonic-boom signatures that have prevented over-land flight. Materials, propulsion, and flight control technologies have developed to the point where, if over-land flight were made possible, a commercial supersonic transport could be economically viable. Computational fluid dynamics, and modern optimization techniques enable designers to reduce the boom signature of candidate aircraft configurations to acceptable levels. However, propulsion systems must be carefully integrated with these low-boom configurations in order that the signatures remain acceptable. One technique to minimize the downward propagation of waves is to mount the propulsion systems above the wing, such that the wing provides shielding from shock waves generated by the inlet and nacelle. This topmounted approach introduces a number of issues with inlet design and performance especially with the highly-swept wing configurations common to low-boom designs. A 1.79%-scale aircraft model was built and tested at the NASA Glenn Research Center's 8-by 6-Foot Supersonic Wind Tunnel (8x6 SWT) to validate the configuration's sonic boom signature. In order to evaluate performance of the top-mounted inlets, the starboard flow-through nacelle on the aerodynamic model was replaced by a 2.3%-scale operational inlet model. This integrated configuration was tested at the 8x6 SWT from Mach 0.25 to 1.8 over a wide range of angles-of-attack and yaw. The inlet was also tested in an isolated configuration over a smaller range of angles-of-attack and yaw. A number of boundary-layer bleed configurations were investigated and found to provide a substantial positive impact on pressure recovery and distortion. Installed inlet performance in terms of mass capture, pressure recovery, and distortion over the Mach number range at the design angle-of-attack of 4-degrees is presented herein and compared to that at 0- degrees, as well as the isolated inlet configuration to highlight installation effects. Performance of the installed inlet fell below that of the isolated inlet at Mach numbers of 1.4 and greater. The installed inlet demonstrated adequate operability over the expected range of angles-of-attack and yaw, but did exhibit definite angle-ofattack and yaw limits at supersonic conditions. At each supersonic flight Mach number, performance parameters near zero yaw angle were relatively insensitive to yaw, but in general the yaw angle yielding best performance was non-zero and varied with angle-of-attack. Performance of the installed inlet is also presented as functions of angle-of-attack and yaw to highlight these effects. Distortion at the aerodynamic interface plane ranged between 10 and 25% at the inlet critical points over the range of flight Mach numbers tested and did not decrease significantly for the isolated inlet. Although these distortion levels would be considered high for operation with a turbine engine, the over-wing installation is likely not as significant a contributor as the low test Reynolds number. This is demonstrated by comparing CFD analysis of the isolated inlet at test scale with that at intermediate and full scales.

Performance↗
NASA NTRS2012

DDS-Suite - A Dynamic Data Acquisition, Processing, and Analysis System for Wind Tunnel Testing

Wind Tunnels have optimized their steady-state data systems for acquisition and analysis and even implemented large dynamic-data acquisition systems, however development of near real-time processing and analysis tools for dynamic-data have lagged. DDS-Suite is a set of tools used to acquire, process, and analyze large amounts of dynamic data. Each phase of the testing process: acquisition, processing, and analysis are handled by separate components so that bottlenecks in one phase of the process do not affect the other, leading to a robust system. DDS-Suite is capable of acquiring 672 channels of dynamic data at rate of 275 MB / s. More than 300 channels of the system use 24-bit analog-to-digital cards and are capable of producing data with less than 0.01 of phase difference at 1 kHz. System architecture, design philosophy, and examples of use during NASA Constellation and Fundamental Aerodynamic tests are discussed.

Burnside, Jathan J.↗
NASA NTRS1998

Combined Numerical/Analytical Perturbation Solutions of the Navier-Stokes Equations for Aerodynamic Ejector/Mixer Nozzle Flows

In spite of rapid advances in both scalar and parallel computational tools, the large number of variables involved in both design and inverse problems make the use of sophisticated fluid flow models impractical, With this restriction, it is concluded that an important family of methods for mathematical/computational development are reduced or approximate fluid flow models. In this study a combined perturbation/numerical modeling methodology is developed which provides a rigorously derived family of solutions. The mathematical model is computationally more efficient than classical boundary layer but provides important two-dimensional information not available using quasi-1-d approaches. An additional strength of the current methodology is its ability to locally predict static pressure fields in a manner analogous to more sophisticated parabolized Navier Stokes (PNS) formulations. To resolve singular behavior, the model utilizes classical analytical solution techniques. Hence, analytical methods have been combined with efficient numerical methods to yield an efficient hybrid fluid flow model. In particular, the main objective of this research has been to develop a system of analytical and numerical ejector/mixer nozzle models, which require minimal empirical input. A computer code, DREA Differential Reduced Ejector/mixer Analysis has been developed with the ability to run sufficiently fast so that it may be used either as a subroutine or called by an design optimization routine. Models are of direct use to the High Speed Civil Transport Program (a joint government/industry project seeking to develop an economically.viable U.S. commercial supersonic transport vehicle) and are currently being adopted by both NASA and industry. Experimental validation of these models is provided by comparison to results obtained from open literature and Limited Exclusive Right Distribution (LERD) sources, as well as dedicated experiments performed at Texas A&M. These experiments have been performed using a hydraulic/gas flow analog. Results of comparisons of DREA computations with experimental data, which include entrainment, thrust, and local profile information, are overall good. Computational time studies indicate that DREA provides considerably more information at a lower computational cost than contemporary ejector nozzle design models. Finally. physical limitations of the method, deviations from experimental data, potential improvements and alternative formulations are described. This report represents closure to the NASA Graduate Researchers Program. Versions of the DREA code and a user's guide may be obtained from the NASA Lewis Research Center.

DeChant, Lawrence Justin↗
NASA NTRS1986

Design of 9.271-pressure-ratio 5-stage core compressor and overall performance for first 3 stages

Overall aerodynamic design information is given for all five stages of an axial flow core compressor (74A) having a 9.271 pressure ratio and 29.710 kg/sec flow. For the inlet stage group (first three stages), detailed blade element design information and experimental overall performance are given. At rotor 1 inlet tip speed was 430.291 m/sec, and hub to tip radius ratio was 0.488. A low number of blades per row was achieved by the use of low-aspect-ratio blading of moderate solidity. The high reaction stages have about equal energy addition. Radial energy varied to give constant total pressure at the rotor exit. The blade element profile and shock losses and the incidence and deviation angles were based on relevant experimental data. Blade shapes are mostly double circular arc. Analysis by a three-dimensional Euler code verified the experimentally measured high flow at design speed and IGV-stator setting angles. An optimization code gave an optimal IGV-stator reset schedule for higher measured efficiency at all speeds.

Steinke, Ronald J.↗
NASA NTRS1997

Combined Numerical/Analytical Perturbation Solutions of the Navier-Stokes Equations for Aerodynamic (Ejector Nozzle) Flows

In spite of the rapid advances in both scalar and parallel computational tools, the large number and breadth of variables involved in aerodynamic systems make the use of parabolized or even boundary layer fluid flow models impractical for both preliminary design and inverse design problems. Given this restriction, we have concluded that reduced or approximate models are an important family of tools for design purposes. This study of a combined perturbation/numerical modeling methodology with an application to ejector-mixer nozzles (shown schematically in the following figure) is nearing completion. The work is being funded by a grant from the NASA Lewis Research Center to Texas A&M University. These ejector-mixer nozzle models are designed to be of use to the High Speed Civil Transport Program and may be adopted by both NASA and industry. A computer code incorporating the ejector-mixer models is under development. This code, the Differential Reduced Ejector/Mixer Analysis (DREA), can be run fast enough to be used as a subroutine or to be called by a design optimization routine. Simplified conservation equations--x-momentum, energy, and mass conservation--are used to define the model. Unlike other preliminary design models, DREA requires minimal empirical input and includes vortical mixing and a fully compressible formulation among other features. DREA is being validated by comparing it with results obtained from open literature and proprietary industry data. Preliminary results for a subsonic ejector and a supersonic ejector are shown. In addition, dedicated experiments have been performed at Texas A&M. These experiments use a hydraulic/gas flow analog to provide information about the inviscid mixing interface structure. Final validation and documentation of this work is expected by May of 1997. However, preliminary versions of DREA can be expected in early 1997. In summary, DREA provides a sufficiently detailed and realistic ejector-mixer nozzle model at a computational cost compatible with preliminary design applications.

DeChant, Lawrence J.↗
NASA NTRS1992

Automatic differentiation of advanced CFD codes for multidisciplinary design

Automated multidisciplinary design of aircraft and other flight vehicles requires the optimization of complex performance objectives with respect to a number of design parameters and constraints. The effect of these independent design variables on the system performance criteria can be quantified in terms of sensitivity derivatives which must be calculated and propagated by the individual discipline simulation codes. Typical advanced CFD analysis codes do not provide such derivatives as part of a flow solution; these derivatives are very expensive to obtain by divided (finite) differences from perturbed solutions. It is shown that sensitivity derivatives can be obtained accurately and efficiently using the ADIFOR source translator for automatic differentiation. In particular, it is demonstrated that the 3-D, thin-layer Navier-Stokes, multigrid flow solver called TLNS3D is amenable to automatic differentiation in the forward mode even with its implicit iterative solution algorithm and complex turbulence modeling. It is significant that by using computational differentiation, consistent discrete nongeometric sensitivity derivatives have been obtained from an aerodynamic 3-D CFD code in a relatively short time, e.g., O(man-week) not O(man-year).

Bischof, C.↗
NASA NTRS1974

Wind tunnel tests of a symmetrical airfoil with scoop fed slots

The design and wind tunnel test of a model vertical tail fin is described in this report. The model is designed to provide the aerodynamic forces necessary for lateral stabilization without moving parts or a separate source of power. It employs scoop-fed slots on both surfaces of the symmetrical airfoil. They are to be controlled differentially by means of a fluidic amplifier to implement an automatic fulltime lateral stabilization system. The results of tests show that the control of forces is stable and quite linear in various modes of operation. Significant forces were produced that can be increased as necessary by increasing slot size and scoop size. Slots can be located ahead of the conventional rudder and the scoop can be at the base of the vertical tail fin to avoid the need for major changes in conventional aircraft design. The first phase of the work demonstrated the feasibility of no-moving-parts aircraft control. The second phase established that a practical fluidic amplifier can be built to control slot flows from fluidic signals. Recommendations are made to optimize the design of the fluidic amplifier and to characterize its dynamic response in support of further analytical studies.

Belsterling, C. A.↗
NASA NTRS2010

Design and Implementation of the Boundary Layer Transition Flight Experiment on Space Shuttle Discovery

In an effort to better the understanding of high speed aerodynamics, a series of flight experiments were installed on Space Shuttle Discovery during the STS-119 and STS-128 missions. This experiment, known as the Boundary Layer Transition Flight Experiment (BLTFE), provided the technical community with actual entry flight data from a known height protuberance at Mach numbers at and above Mach 15. Any such data above Mach 15 is irreproducible in a laboratory setting. Years of effort have been invested in obtaining this valuable data, and many obstacles had to be overcome in order to ensure the success of implementing an Orbiter modification. Many Space Shuttle systems were involved in the installation of appropriate components that revealed 'concurrent engineering' was a key integration tool. This allowed the coordination of all various parts and pieces which had to be sequenced appropriately and installed at the right time. Several issues encountered include Orbiter configuration and access, design requirements versus current layout, implementing the modification versus typical processing timelines, and optimizing the engineering design cycles and changes. Open lines of communication within the entire modification team were essential to project success as the team was spread out across the United States, from NASA Kennedy Space Center in Florida, to NASA Johnson Space Center in Texas, to Boeing Huntington Beach, California among others. The forum permits the discussion of processing concerns from the design phase to the implementation phase, which eventually saw the successful flights and data acquisition on STS-119 in March 2009 and on STS-128 in September 2009.

Spanos, Theodoros A.↗
NASA NTRS1993

Controls-structures-interaction dynamics during RCS control of the Orbiter/SRMS/SSF configuration

During the assembly flights of the Space Station Freedom (SSF), the Orbiter will either dock with the SSF and retract to the final berthed position, or will grapple the SSF using the Shuttle Remote Manipulator System (SRMS) and maneuver the SRMS coupled vehicles to their final berthed position. The SRMS method is expected to take approximately one to one and a half hours to complete and require periodic attitude corrections by either the Orbiter or the SSF reaction control system (RCS) or continuous control by a control moment gyro (CMG) system with RCS desaturation as required. Free drift of the attached vehicles is not currently thought to be acceptable because the desired system attitude will quickly deteriorate due to unbalanced gravity gradient and aerodynamic torques resulting in power generation problems, thermodynamic control problems, and communications problems. This paper deals with the simulation and control of the SRMS during trunnion/latch interaction dynamics and during RCS maneuvers. The SRMS servo drive joints have highly non-linear elastic characteristics which tend to degrade sensitive control strategies. In addition the system natural frequencies are extremely low and depend on the drive joint deflections and SRMS geometric position. The lowest mean period of oscillation for the Orbiter/SRMS/SSF(MB6) system in brakes hold mode positioned near the final berthed position is approximately 120 seconds. A detailed finite element model of the SRMS has been developed and used in a newly developed SRMS systems dynamics simulation to investigate the non-linear transient response dynamics of the Orbiter/SRMS/SSF systems. The present SRMS control strategy of brakes only recommended by the Charles Draper Labs is contrasted with a robust controller developed by the authors. The robust controller uses an optimal inear quadratic regulator (LQR) to optimally place the closed-loop poles of a multivariable continuous-time system within the common region of an open sector with the sector angle plus or minus 45 degrees from the negative real axis, and the left-hand side of a parallel to the imaginary axis in the complex s-plane. This guarantees that the critical damping ratio for the desired control modes is equal to or in excess of 0.707. The matrix sign function is used for solving the Riccati equations which appear in the controller design procedure. Fast and stable algorithms have recently been developed for the computation of the matrix sign function. Simulation results are given which demonstrate the potential CSI involvement for the current SRMS control system and the proposed control system.

Schliesing, J. A.↗
NASA NTRS1999

Overview of Current Turbine Aerodynamic Analysis and Testing at MSFC

An overview of the current turbine aerodynamic analysis and testing activities at NASA/Marshall Space Flight Center (MSFC) is presented. The presentation is divided into three areas. The first area is the three-dimensional (3D), unsteady Computational Fluid Dynamics (CFD) analysis of the Fastrac turbine. Results from a coupled nozzle, blade, and exit guide vane analysis and from an uncoupled nozzle and coupled blade and exit guide vane will be presented. Unsteady pressure distributions, frequencies, and exit profiles from each analysis will be compared and contrasted. The second area is the testing and analysis of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) turbine with instrumented first stage blades. The SSME HPFTP turbine was tested in air at the MSFC Turbine Test Equipment (TTE). Pressure transducers were mounted on the first stage blades. Unsteady, 3D CFD analysis was performed for this geometry and flow conditions. A sampling of the results will be shown. The third area is a status of the Turbine Performance Optimization task. The objective of this task is to improve the efficiency of a turbine for potential use on a next generation launch vehicle. This task includes global optimization for the preliminary design, detailed optimization for blade shapes and spacing, and application of advanced CFD analysis. The final design will be tested in the MSFC TTE.

Griffin, Lisa W.↗
NASA NTRS1980

Optimum subsonic, high-angle-of-attack nacelles

Many proposed advanced aircraft - but especially tilt-nacelle, subsonic-cruise, V/STOL aircraft - require nacelles that operate over a wide range of aerodynamic conditions. The optimum design of such nacelles and their inlets is described, including how the inlet low-speed design conditions are selected, the conditions for which the various regions of the inlet are designed, and appropriate criteria of merit. For low-speed operation the optimum internal surface velocity distributions and skin friction distributions are described for three categories of inlets: those without boundary-layer control (BLC), those with BLC, and those with blow-in door slots and retractable slats. Experimental results are presented that show the performance of the various types of inlets. At cruise speed the effect of factors that reduce the nacelle external surface area and the local skin friction is illustrated. These factors are cruise Mach number, inlet throat size, fan-face Mach number, and nacelle contour. The interrelation of these cruise-speed factors with the design requirements for good low-speed performance is discussed. Finally an inlet design without BLC and an optimized inlet design with slots and slats are compared to illustrate the possible reductions in nacelle size.

Luidens, R. W.↗
NASA NTRS2014

Regional Characteristics of the Relationship Between Columnar AOD and Surface PM2.5: Application of Lidar Aerosol Extinction Profiles over Baltimore-Washington Corridor During DISCOVER-AQ

The first field campaign of DISCOVER-AQ (Deriving Information on Surface conditions from COlumn and VERtically resolved observations relevant to Air Quality) took place in July 2011 over Baltimore-Washington Corridor (BWC). A suite of airborne remote sensing and in-situ sensors was deployed along with ground networks for mapping vertical and horizontal distribution of aerosols. Previous researches were based on a single lidar station because of the lack of regional coverage. This study uses the unique airborne HSRL (High Spectral Resolution Lidar) data to baseline PM2.5 (particulate matter of aerodynamic diameter less than 2.5 μm) estimates and applies to regional air quality with satellite AOD (Aerosol Optical Depth) retrievals over BWC (∼6500 sq. km). The linear approximation takes into account aerosols aloft above AML (Aerosol Mixing Layer) by normalizing AOD with haze layer height (i.e., AOD/HLH). The estimated PM2.5 mass concentrations by HSRL AOD/HLH are shown within 2 RMSE (Root Mean Square Error ∼9.6 μg/cu. m) with correlation ∼0.88 with the observed over BWC. Similar statistics are shown when applying HLH data from a single location over the distance of 100 km. In other words, a single lidar is feasible to cover the range of 100 km with expected uncertainties. The employment of MPLNET-AERONET (MicroPulse Lidar NETwork - AErosol RObotic NETwork) measurements at NASA GSFC produces similar statistics of PM2.5 estimates as those derived by HSRL. The synergy of active and passive remote sensing aerosol measurements provides the foundation for satellite application of air quality on a daily basis. For the optimal range of 10 km, the MODIS-estimated PM2.5 values are found satisfactory at 27 (out of 36) sunphotometer locations with mean RMSE of 1.6-3.3 μg/cu. m relative to PM2.5 estimated by sunphotometers. The remaining 6 of 8 marginal sites are found in the coastal zone, for which associated large RMSE values ∼4.5-7.8 μg/cu. m are most likely due to overestimated AOD because of water-contaminated pixels.

DISCOVER-AQ↗
NASA NTRS2012

Characterization of Aerodynamic Performance of Boundary-Layer-Ingesting Inlet Under Crosswind

NASA has been studying future transport concepts, envisioned to be technically realizable in the timeframe of 2020-2030, to meet environmental and performance goals. One concept receiving considerable interest involves a propulsion system embedded into a hybrid wing-body aircraft. While offering significant advantages in fuel savings and noise reduction by this concept, there are several technical challenges that are not encountered in the current fleet and must be overcome so as to deliver target performance and operability. One of these challenges is associated with an inlet system that ingests a significantly thick boundary layer, developing along the wing-body surface, into a serpentine diffuser before the flow meeting fan blades. The flow is subject to considerable total pressure loss and distorted at the fan face, much more significantly than in the inlet system of conventional aircraft. In our previous studies [1, 2], we have shown that through innovative design changes on the airframe surface, it is possible to simultaneously increase total pressure recovery and decrease distortion in the flow, without resorting to conventional penalty-ridden flow control concepts, such as vortex generator or boundary layer bleeding/suction. In the current study, we are interested in understanding the following issues: how the embedded propulsion system performs under a crosswind condition by studying in detail the flow characteristics of two inlets, the baseline and another optimized previously under the cruise condition. With the insight, it is hoped that it can help in the follow-on study by devising effective strategies to minimize flow distortion arising from the integration of an embedded-engine system into an airframe to the level acceptable to the operation and fuel consumption before 2030. To achieve these demanding goals, non-conventional concepts are called for; but technology gap is too big that it requires evolutionary approach by focusing various concepts and technologies needed in the next three generations of aircraft, respectively named as N+1, N+2, and N+3. Noticeably, considerable reduction in each category of 1 is required in N+2 (relative to Boeing 777-200 and GE90 engines) and N+3 (relative to Boeing 737-800 and CFM56-7B engines). In this study, concepts for N+2 is our interest. A concept that has potential to achieve these metrics and has been under intensive study is the hybrid wing body (HWB) airframe with a tightly integrated propulsion system, see Fig. 1. The inlet is non-circular at the entrance and the entering flow, no longer uniform or free of disturbances, and is now carrying with it a boundary layer developing along the fuselage; the inlet is thus known as boundary-layer-ingesting (BLI) inlet.

Liou, Meng-Sing↗
NASA NTRS2023

Advances in Aero-Propulsive Modeling for Fixed-Wing and eVTOL Aircraft Using Experimental Data

Small unmanned aircraft and electric vertical takeoff and landing (eVTOL) aircraft have recently emerged as vehicles able to perform new missions and stimulate future air transportation methods. This dissertation presents several system identification research advancements for these modern aircraft configurations enabling accurate mathematical model development for flight dynamics simulations based on wind-tunnel and flight-test data. The first part of the dissertation focuses on advances in flight-test system identification methods using small, fixed-wing, remotely-piloted, electric, propeller-driven aircraft. A generalized approach for flight dynamics model development for small fixed-wing aircraft from flight data is described and is followed by presentation of novel flight-test system identification applications, including: aero-propulsive model development for propeller aircraft and nonlinear dynamic model identification without mass properties. The second part of the dissertation builds on established fixed-wing and rotary-wing aircraft system identification methods to develop modeling strategies for transitioning, distributed propulsion, eVTOL aircraft. Novel wind-tunnel experiment designs and aero-propulsive modeling approaches are developed using a subscale, tandem tilt-wing, eVTOL aircraft, leveraging design of experiments and response surface methodology techniques. Additionally, a method applying orthogonal phase-optimized multisine input excitations to aircraft control effectors in wind-tunnel testing is developed to improve test efficiency and identified model utility. Finally, the culmination of this dissertation is synthesis of the techniques described throughout the document to form a flight-test system identification approach for eVTOL aircraft that is demonstrated using a high-fidelity flight dynamics simulation. The research findings highlighted throughout the dissertation constitute substantial progress in efficient empirical aircraft modeling strategies that are applicable to many current and future aeronautical vehicles enabling accurate flight simulation development, which can subsequently be used to foster advancement in many other pertinent technology areas.

system identification↗
NASA NTRS1984

Fuel conservative guidance concept for shipboard landing of powered-life aircraft

A simulation study was undertaken to investigate the application of energy conservative guidance (ECG) software, developed at NASA Ames Research Center, to improve the time and fuel efficiency of powered lift airplanes operating from aircraft carriers at sea. When a flightpath is indicated by a set of initial conditions for the aircraft and a set of positional waypoints with associated airspeeds, the ECG software synthesizes the necessary guidance commands to optimize fuel and time along the specified path. A major feature of the ECG system is the ability to synthesize a trajectory that will allow the aircraft to capture the specified path at any waypoint with the desired heading and airspeed from an arbitrary set of initial conditions. Five paths were identified and studied. These paths demonstrate the ECG system's ability to save flight time and fuel by more efficiently managing the aircraft's capabilities. Results of this simulation study show that when restrictions on the approach flightpath imposed for manual operation are removed completely, fuel consumption during the approach was reduced by as much as 49% (610 lb fuel) and the time required to fly the flightpath was reduced by as much as 41% (5 min). Savings due to ECG were produced by: (1) shortening the total flight time; (2) keeping the airspeed high as long as possible to minimize time spent flying in a regime in which more engine thrust is required for lift to aid the aerodynamic lift; (3) minimizing time spent flying at constant altitude at slow airspeeds; and (4) synthesizing a path from any location for a direct approach to landing without entering a holding pattern or other fixed approach path.

Warner, D. N., Jr.↗
NASA NTRS2014

A Modular Aerospike Engine Design Using Additive Manufacturing

A modular aerospike engine concept has been developed with the objective of demonstrating the viability of the aerospike design using additive manufacturing techniques. The aerospike system is a self-compensating design that allows for optimal performance over the entire flight regime and allows for the lowest possible mass vehicle designs. At low altitudes, improvements in Isp can be traded against chamber pressure, staging, and payload. In upper stage applications, expansion ratio and engine envelope can be traded against nozzle efficiency. These features provide flexibility to the System Designer optimizing a complete vehicle stage. The aerospike concept is a good example of a component that has demonstrated improved performance capability, but traditionally has manufacturing requirements that are too expensive and complex to use in a production vehicle. In recent years, additive manufacturing has emerged as a potential method for improving the speed and cost of building geometrically complex components in rocket engines. It offers a reduction in tooling overhead and significant improvements in the integration of the designer and manufacturing method. In addition, the modularity of the engine design provides the ability to perform full scale testing on the combustion devices outside of the full engine configuration. The proposed design uses a hydrocarbon based gas-generator cycle, with plans to take advantage of existing powerhead hardware while focusing DDT&E resources on manufacturing and sub-system testing of the combustion devices. The major risks for the modular aerospike concept lie in the performance of the propellant feed system, the structural integrity of the additive manufactured components, and the aerodynamic efficiency of the exhaust flow.

Peugeot, John↗
NASA NTRS1993

Aerodynamic control of NASP-type vehicles through Vortex manipulation. Volume 2: Static wind tunnel tests

Forebody Vortex Control (FVC) was explored in this research program for potential application to a NASP-type configuration. Wind tunnel tests were conducted to evaluate a number of jet blowing schemes. The configuration tested has a slender forebody and a 78 deg swept delta wing. Blowing jets were implemented on the leeward side of the forebody with small circular tubes tangential to the surface that could be directed aft, forward, or at angles in between. The effects of blowing are observed primarily in the yawing and rolling moments and are highly dependent on the jet configuration and the angle of attack. Results show that the baseline flow field, without blowing activated, is quite sensitive to the geometry differences of the various protruding jets, as well as being sensitive to the blowing, particularly in the angle of attack range where the forebody vortices are naturally asymmetric. The time lag of the flow field response to the initiation of blowing was also measured. The time response was very short, on the order of the time required for the flow disturbance to travel the distance from the nozzle to the specific airframe location of interest at the free stream velocity. Overall, results indicate that sizable yawing and rolling moments can be induced with modest blowing levels. However, direct application of this technique on a very slender forebody would require thorough wind tunnel testing to optimize the jet location and configuration.

Suarez, Carlos J.↗