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进气道2

2011-09-15 24页 pdf 2MB 24阅读

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进气道2 1 American Institute of Aeronautics and Astronautics Distribution A: Approved for public release; distribution unlimited: Case Number 88ABW-2010-4076 A Reusable, Rocket and Airbreathing Combined Cycle Hypersonic Vehicle Design for Access-to-Space Ajay P....
进气道2
1 American Institute of Aeronautics and Astronautics Distribution A: Approved for public release; distribution unlimited: Case Number 88ABW-2010-4076 A Reusable, Rocket and Airbreathing Combined Cycle Hypersonic Vehicle Design for Access-to-Space Ajay P. Kothari 1 Astrox Corporation, College Park, MD 20740 John W. Livingston 2 Aerospace Systems Design and Analysis, ASC/XRE, Wright Patterson AFB, OH 45433 Christopher Tarpley 3 Astrox Corporation, Colorado Springs, CO 80919 Venkatraman Raghavan 4 Astrox Corporation, College Park, MD 20740 and Kevin G. Bowcutt 5 and Thomas R. Smith 6 Boeing Corporation, Huntington Beach, CA 92647 Several rocket boosted RBCC (Rocket-Based Combined Cycle) designs were created and analyzed for suitability for an access-to-space TSTO vehicle system. Level 1 analysis was performed that included vehicle closure, weight breakdown (to component level), flight trajectory data, propulsion and aerodynamic performance. Various inlet shapes were considered, such as symmetric and non-symmetric inlets. Also, several geometric configurations were studied, such as single flowpath, dual flowpath, engine-on-top and engine-on-bottom. At the end of the study, the number of candidate designs were reduced to two; one as a primary design and the other as the backup design. The primary design, a dual flowpath, engine-on-top design was selected for further analysis. Due to the large volume of the payload, the dual flowpath design was found to be more suitable than a single flowpath design. Nomenclature A_capture = flow capture area (frontal projection) A_exit = nozzle exit area (frontal projection) A_front = total frontal area Aeroloss = losses due to drag AoA = angle of attack CAD = computer aided design CR = contraction ratio D = drag DeltaV = measure of energy differential in velocity units 1 President, Astrox HQ, 3500 Marlbrough Way, Ste 100, AIAA Senior Life Member 2 Systems Design Engineer, ASC/XRE, AIAA Member 3 Principal Research Scientist, Astrox, 3310 Van Teylingen Drive, AIAA Senior Member 4 Principal Research Engineer, Astrox HQ, 3500 Marlbrough Way, Ste 100 5 Sr. Technical Fellow and Chief Scientist of Hypersonics, MS H45N-E406, AIAA Fellow 6 Associate Technical Fellow, MS H45N-E408, AIAA Member AIAA SPACE 2010 Conference & Exposition 30 August - 2 September 2010, Anaheim, California AIAA 2010-8905 Copyright © 2010 by Astrox Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. 2 American Institute of Aeronautics and Astronautics DM = dry margin EISP = effective ISP g = acceleration due to gravity Gamma, = flight path angle GNC = guidance and control Gravloss = losses due to gravity HRSI = high-temperature reusable surface insulation (TPS tile) HTHL = horizontal takeoff and horizontal landing Ieff = effective ISP ISP = specific impulse JP-2 = jet fuel kPa = kilo Pascals (pressure in SI units) L = lift LEO = low earth orbit L/H = inlet width to height ratio LHC = liquid hydrocarbon (fuel) LH2 = liquid hydrogen (fuel) LOX = liquid oxygen (oxidizer) MdotF = fuel mass flow rate q = dynamic pressure RBCC = rocket-based combined cycle Rc/H = inlet corner ratio RDP = inlet radial deviation parameter RP-2 = hydrocarbon rocket fuel RR = reusable rocket SSTO = single-stage-to-orbit T = thrust TJ = turbo-jet TBCC = turbine-based combined cycle TPS = thermal protection system TSTO = two-stage-to-orbit VTHL = vertical takeoff and horizontal landing V = freestream velocity W = vehicle weight α = angle of attack I. Introduction N 2008, the Office of the Chief Scientist of the Air Force tasked Astrox Corporation with a study comparing eight Two-Stage-to-Orbit (TSTO) options for placing a 20,000 lbs payload module in Low Earth Orbit (LEO). The options included all rocket (expendable, hybrid and reusable) as well as a combination of reusable rocket and air- breathing propulsion stages including turbine and/or ram/scramjet propulsion. This study was done at Level 0 to Level 1 fidelity. Of the eight options, three showed promise and were selected for further analysis at Level 1 and then at Level 2. NASA elected to pursue the Turbine-Based-Combined-Cycle (TBCC) first stage and Reusable Rocket second stage (TBCC/RR) option. The Air Force elected to pursue the Reusable Rocket first stage and Reusable Rocket second stage (RR/RR) option and Reusable Rocket first stage but Rocket-Based-Combined-Cycle (RBCC) second stage (RR/RBCC) option. Astrox Corporation was the Air Force‟s industry team lead for the RR/RBCC option with Mr. John Livingston being the Government lead. The industry team included Astrox Corporation, Boeing Corporation (Huntington Beach) and GoHypersonics Inc for the CFD part which is not presented herein. The team looked at single flowpath and dual flowpath designs, both with the engines facing the ground (engines-on- bottom) or facing skyward (engines-on-top) while in air-breathing mode. The inlets are inward turning. The dual flowpath design with the engines-on-top was selected from among the four mentioned. The RBCC stages at Mach 3.5. The empty weight of the RBCC is 117K lbs including the dry margin. The system lifts of vertically using first stage rockets but both stages land horizontally. Since the first stage goes only up to Mach 3.5, it will be able to glide I 3 American Institute of Aeronautics and Astronautics back and will require minimal to no TPS. The RBCC will stage at lower Q and dip down to join Q=2000 psf trajectory during Ram/Scram. The orbit insertion rockets using LH2/LOX will take over at Mach 10. Since none of the stages is designed for horizontal take off, the reduction in the wing and gears weight has a cumulatively beneficial effect. Additionally since the thrust to weight of the rocket is much larger than that of turbine, that aspect also has a much beneficial effect. All of these benefits coalesce to yield a very attractive system compared to what has been pursued before which can lead to a paradigm shift in our thinking. The design will have more than 1500 nm cross-range capability and such an RBCC can also be designed to have 500-1000 nm offset capability. II. Background In earlier studies several SSTO and TSTO involving air-breathers have been analyzed 1-7 . Figure 1 identifies some these configurations and their baseline gross weights (the STS and XB-70 Valkyrie are included for scale reference). As seen from the figure, the study investigated eighteen vehicle configurations; nine SSTO and nine TSTO. All of the SSTO vehicles were hypersonic air-breathing vehicles differing by inlet type, propellant selection, low-speed propulsion cycle, and takeoff mode. The TSTO configurations included three pure rocket systems as well as air-breathing vehicles combined with either an upper stage rocket orbiter or first stage rocket booster. The air- breathing vehicles used either an inward-turning “IN” inlet or more traditional wedge “2D” type inlet geometry. The low-speed propulsion cycles for all air-breathers was provided by either integrated rockets “-RB” or turbines “- TB” operating on hydrogen “(H2)” or hydrocarbon “(HC)” fuel. One HTHL vehicle uses a pure turbine booster as a first stage. The TSTO notation in the figure is listed as Stage1 / Stage2. Figure 1. Vehicle Size Comparison 4 American Institute of Aeronautics and Astronautics Figure 2. Empty Weight Comparison (Klbs) Empty Weight Figure 2 shows the empty weight comparison of the SSTO and TSTO vehicles. At this level of analysis, the total vehicle system empty weight may be successfully employed as the main cost driver of a launch vehicle system. Most of the launch operation and flight refurbishment costs, as well as the initial design and procurement costs of a launch vehicle scale roughly with empty weight 8 . When comparing the empty weights as a rough measure of the approximate cost and feasibility of designing and constructing the vehicle 9 it must be remembered that, “pound for pound,” a pure rocket vehicle is likely to cost less than a more advanced and technologically uncertain hypersonic air-breathing vehicle. 5 American Institute of Aeronautics and Astronautics Figure 3. Gross Weight Comparison (Klbs) Gross Weight Vehicle gross takeoff weight as shown in Fig. 3 is often cited as a principle metric of comparison between different vehicle configurations. However, the vehicle gross weight is not as useful a figure of merit as the three listed above. The major constituents of the gross weight for the vehicles are the propellants required. Compared with the cost of acquiring, and maintaining the vehicle, the cost of purchasing each flight‟s propellant is nearly insignificant. While a higher gross weight vehicle for a given mission may represent a lower performing propulsion system, it is the impact of that performance on the vehicle‟s empty weight and surface area that are of the most interest. However, the gross weight was included in this study because it does give quick insight into the scaling of parameters that have to do with the fueled vehicle such as propulsion thrust requirement, pad limitations and in the case of horizontally launched vehicles, the wing and landing gear sizing. It is also used to determine if a particular HTHL vehicle baseline or grown solution has exceeded the runway bearing load limitation which, for this study, was assumed to be 1.5 million lbs. 6 American Institute of Aeronautics and Astronautics Growth Factor The growth factor shown in Fig. 4 is a measure of the scaling response in vehicle empty weight to an increase in the unit weight of the vehicle structure. This increase is often due to a change in the estimation of the corresponding weight of some structural technology; such has a heavier or lighter weight TPS tile type. The growth factor can therefore be used as a measure of the vehicle‟s response to the technological uncertainty inherent in the development of a future system. A general vehicle scaling reaction to such an increase follows the steps outlined below: 1) A closed vehicle solution experiences an increase in a structural unit weight. 2) That percentage increase multiplied by the total amount of that structural component present in the closed solution results in an additional amount of empty weight that must now be carried by the vehicle. 3) The vehicle solution is no longer closed. The additional weight is a perturbing influence that triggers a scaling up of the vehicle solution in order to re-close the vehicle. 4) As the vehicle grows in response to the weight change it must now do so with a correspondingly higher unit weight. 5) This double impact causes a much larger change in the re-closed vehicle‟s empty weight versus the original empty weight than just the addition of the perturbing change in structure weight. 6) The growth factor is obtained by differentiating the empty weight scaling equation 1,5 with respect to the weight change and is the slope of the delta weight / delta weight curve at the point of the vehicle solution. Figure 4. Growth Factors 7 American Institute of Aeronautics and Astronautics III. Design Requirements The primary focus of this work is the design and analysis, to Level 1, of the Air Force‟s RR/RBCC option. This is a TSTO system consisting of a rocket powered booster stage with an RBCC upper stage carrying the payload to low earth orbit. The booster stage is a reusable rocket that carries the upper stage to Mach 3.5 at an altitude of 73,918 ft where the dynamic pressure is 696 psf. This allows for a low q staging of the upper stage. The upper stage is a dual flowpath RBCC design that flies using a ram/scramjet to Mach 10.0. The RBCC stage transitions to rocket and completes the ascent to orbit. This system takes off vertically and both stages land horizontally. The design requirements can be summarized as follows: 1) Level 1 Design 2) Reusable Rocket Booster (LHC fuel) 3) Ram-Scram Upperstage (Methane fuel) 4) Upperstage uses LH2 rockets for ascent to orbit 5) Payload weight 20,000 lbs, payload size is 12‟ diameter and 30‟ length, payload volume is 3393 ft3 6) Landing Speed not greater than SS Orbiter viz. 188 knots Choosing a vertical launch system using an RBCC upper stage has several advantages. As seen in Fig. 5, the RR/RBCC is the smallest of the reusable systems. The pros and cons of an RBCC system are enumerated below. Advantages: 1) Substantially reduced gear weights and no turbine weight or volume issues (compared to TBCC) 2) Rocket weights are substantially less than turbine weights (compared to TBCC) 3) No turbine flowpath integration issues (compared to TBCC) 4) Smaller wing size (compared to HTHL options) 5) Can be sized up for larger payload or more constraining orbits like polar or GEO transfer (compared to HTHL) 6) Has the possibility to extend to SSTO Disadvantages: 1) Reentry TPS needed. Hence higher TPS weights and more heat constrained (compared to TBCC and RR). 2) More complexities (compared to RR) IV. Two-Stage RBCC Design A. Methodology Used 1) Define fidelity levels applicable to Reusable Rocket, RBCC and TBCC designs 2) Create a single flowpath design, finding suitable leading edge shapes 3) Re-Entry 3DOF trajectory/coding, data interface development, reentry POST3D phase creation, testing and verification of result 4) External TPS sizing/research into methodology 5) Upgrade all components to Level 1 6) Create necessary vehicles and vehicle systems 8 American Institute of Aeronautics and Astronautics Figure 5. Comparison of TSTO Vehicles Figure 5 shows a comparison of various RR, RBCC, and TBCC options. As can be seen, the RBCC system, Option 3, is substantially smaller than all but the expendable launch system shown as Option 1. B. Assumptions, and Procedures The HySIDE© software was used to build the RBCC upper stage system, and to resize the vehicle until closed. The methods used in HySIDE software are the following: 1) HySIDE is used to build all the stages, to combine them as TSTO, and to resize each of them till closed properly. 2) Method of Characteristic with Reference Enthalpy method for viscous calculations were used to design and shape the flowpaths and obtain all thermodynamic properties at each mesh point 10,11 . 3) The number mesh points for each stage vehicle ranged from 10,000-15,000. 4) Shock expansion technique was used for the external flow areas such as wings, tails and external surfaces 5) On-design Tip-to-Tail ISP was computed by integrating all areas at all mesh points. 6) Rocket ISP was calculated using the requisite rocket type from among the 30 different rocket models incorporated into HySIDE. 7) Missile DATCOM was used for aerodynamic performance throughout the trajectory flown. 8) Astrox Trajectory code was used for the ascent trajectory (3 DOF). 9) The vehicles were resized until “closed”, starting from the upper stage. 10) Each resizing performs calculations again at all the 10-15 thousand mesh points and reintegrates. 11) The POST3D code was used to study the RBCC reentry calculations. RBCC aerodynamic data was used to perform RBCC reentry calculations. The POST3D calculations and plotting capabilities were added to HySIDE software. C. Design Issues During the design cycle of the TSTO RBCC system, the following issues were addressed: 1) Dual Flowpath vs. Single Flowpath 2) Engine(s) on top or bottom? 9 American Institute of Aeronautics and Astronautics 3) Staging Mach number 4) Fuel for Ram/Scram cycle 5) Propellants for Rocket segments 6) Scram-jet cutoff speed 7) A_exit/A_capture 8) Inlet contraction ratio 9) Isolator and Combustor lengths 10) Re-entry and Payload Deployment Accelerator Area Ratios At the end of the scramjet cycle, not only does the vehicle run out of ISP, or more precisely, EISP (Ieff), but it begins to run out of acceleration, (T-D)/W. One way to increase this is by increasing T, which means it should capture as much air as possible. Basically the vehicle should not be designed like a cruiser. A_front/A_capture should be as close to one as possible, and A_exit/A_capture should also be one, both in order to reduce drag. This however, severely restricts volume available and a compromise is needed. From past experience, this number (A_exit/A_capture) seems to gravitate to 1.8-2.2 for accelerators. Some important frontal projected areas at design point are: 1) A_capture/A_front ~ 0.45 2) A_exit/A_front ~ 0.85 3) A_exit/A_capture ~ 1.9 Fuel selection A denser propellant (average density) for the lower rocket segments (takeoff) is best, even if it offers lower ISP such as hydrocarbon rocket fuel (RP-2) as fuel and liquid oxygen as oxidizer. Fuel to carry smaller structure to orbit is less costly overall than the additional fuel that you are taking up to some Mach number. Why? A higher ISP propellant (e.g. LH2/LOX) during the later phases of flight, even if it means having lower density and more volume. This is especially true for a single stage RBCC solution, but it is also true for a TSTO whether it is rocket/rocket or rocket/RBCC. Methane was selected as the fuel for scram-jet operations. Methane has the advantage of being a soft- cryogen, not voluminous like liquid hydrogen, acceptable Isp and able to perform in scram-jet mode to a higher Mach number than JP fuel. For the ascent to orbit rocket stage, liquid hydrogen was used as the fuel with liquid oxygen as the oxidizer. Re-entry It is of primary importance to prevent hot air from flowing through the inlet and the flowpath during reentry. One can think of trying to close up the inlet but this is cumbersome especially for the inward inlet shapes. One easier and smarter choice would be to reenter on its back. This means that the back will need to have, AT LEAST, the same tile structure that the Space Shuttle does – covered with HRSI on most of the area, and perhaps more severe if the vehicle intends to fly on its back during ascent (Ram/Scram) also. Now keeping in mind the fact that the ascent portion of the trajectory (flying at q=2000 psf for almost 20 minutes) is a more severe heat environment, the ascent will dictate the tile structure there, and this may be sufficient for reentry anyways. This means that the engines will have to be on top (single module) or near the top (dual module). This also means that the inlet and flowpath shape will have to be designed, using streamline tracing, such that the vehicle back would be flat or a waverider, the engine would be on the opposite side, and would be on the top during ascent, descent and landing. Angle of Attack The vehicle wants fly at 3.5-5 deg AoA between Mach 3.5 to 10 (Airbreathing portion). At these times, the engines are designed and placed such that they get full air capture when the vehicle is flying at 4.5 deg AoA measured with respect to the bottom surface. 10 American Institute of Aeronautics and Astronautics Other Considerations The help provided by a strut rocket at low Mach numbers is more th
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