Sponsor Message:
Aviation Technical / Operations Forum
My Starred Topics | Profile | New Topic | Forum Index | Help | Search 
SST Propulsion Emissions And Noise.  
User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Posted (7 years 9 months 2 weeks 10 hours ago) and read 1927 times:

Just a thought, assuming all the criteria one day are met to build an efficient, low-noise, and low sonic-boom SST.

In the light of the current climate change issues regarding long haul flights, would it be better to have a fleet of (perhaps wide-body?) SSTs than a fleet of wide-body subsonic airliners? Recent studies indicate that the effect of NOx is smaller than had been thought in the 1970s. The use of combustors with an emission index of about 5 grammes of NOx produced for every kilogram of fuel burned may be sufficient to keep NOx within acceptable limits. Achieving an index of 5 represents a reduction of 70 to 90% percent relative to current engines operating at similar conditions. NASA's HSR program conducted small-scale steady state and transient tests of combustor rigs to demonstrate that emissions indexes below 5g/kg could be achieved. Unfortunately the HSR program was terminated before an engine with such combustors could be tested in flight.

SSTs cruising well above the stratospheric ozone layer may still create problems because of water vapor emissions in the stratosphere. The stratosphere naturally contains very little water. Water destroys ozone and provides a source of HOx that enhances ozone loss. However, ozone depletion could be eliminated by reducing average cruise altitudes to about 50.000- to 60.000 ft.

I'm aware that the take-off, climb, descent and approach phases maybe less efficient for the SST when compared to a subsonic airliner. On the other hand the SSTs climb performance is better so time to cruise speed and altitude is much shorter.

Taking total flight time into account, of which cruise is the longest (most efficient) segment, would it be fair to say that less pollution takes place flying an SST long distance for 2 to 3 hours than a lumbering subsonic wide-body for 8 to 9 hours?

17 replies: All unread, jump to last
 
User currently offlineStarlionblue From Greenland, joined Feb 2004, 17055 posts, RR: 67
Reply 1, posted (7 years 9 months 2 weeks 9 hours ago) and read 1922 times:

Quoting Starglider (Thread starter):
Taking total flight time into account, of which cruise is the longest (most efficient) segment, would it be fair to say that less pollution takes place flying an SST long distance for 2 to 3 hours than a lumbering subsonic wide-body for 8 to 9 hours?

Maybe. But while we're at it, why don't we just make the thing sub-orbital and only use the engines for takeoff/climb and approach/landing.



"There are no stupid questions, but there are a lot of inquisitive idiots."
User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 2, posted (7 years 9 months 2 weeks 6 hours ago) and read 1911 times:

Quoting Starlionblue (Reply 1):
Maybe. But while we're at it, why don't we just make the thing sub-orbital and only use the engines for takeoff/climb and approach/landing.

Although i agree that would be a favorable concept, it is beyond the scope of technology when compared to materials and powerplants used in an SST.

A next generation SST will most likely use high-temperature composite materials. Selecting the right structures and materials for an SST airframe designed to fly 60,000 hours in its lifetime, in temperatures approaching 350 degrees Fahrenheit, is critical to making a future supersonic transport economically feasible. Weight and manufacturing costs must be minimized, while strength and durability are maintained.

Going sub-orbital makes the craft essentially a space-craft and requires more heat tolerance, exotic materials such as super alloys and refractory alloys. Don't even want to speculate what the life-span and maintenance program for such a vehicle would be. That is setting the goal a step higher than i intend this thread to be.


User currently offlineStarlionblue From Greenland, joined Feb 2004, 17055 posts, RR: 67
Reply 3, posted (7 years 9 months 2 weeks 4 hours ago) and read 1899 times:

Quoting Starglider (Reply 2):
That is setting the goal a step higher than i intend this thread to be.

Fair enough.



"There are no stupid questions, but there are a lot of inquisitive idiots."
User currently offlineB2707SST From United States of America, joined Apr 2003, 1369 posts, RR: 59
Reply 4, posted (7 years 9 months 1 week 6 days 21 hours ago) and read 1883 times:

Quoting Starglider (Thread starter):
In the light of the current climate change issues regarding long haul flights, would it be better to have a fleet of (perhaps wide-body?) SSTs than a fleet of wide-body subsonic airliners? Recent studies indicate that the effect of NOx is smaller than had been thought in the 1970s. The use of combustors with an emission index of about 5 grammes of NOx produced for every kilogram of fuel burned may be sufficient to keep NOx within acceptable limits. Achieving an index of 5 represents a reduction of 70 to 90% percent relative to current engines operating at similar conditions. NASA's HSR program conducted small-scale steady state and transient tests of combustor rigs to demonstrate that emissions indexes below 5g/kg could be achieved. Unfortunately the HSR program was terminated before an engine with such combustors could be tested in flight.

While it's true that enormous advances in NOx reduction have been made, these breakthroughs are as applicable to subsonic turbofans as to supersonic engines. SSTs have gotten more attention because they tend to cruise in the band of maximum ozone density, so their NOx emissions have a much greater impact than those of subsonic jets flying at half the altitude. If NOx emissions become a critical issue, we'll see low-NOx processes like lean premixed prevaporized or rich-burn/quick-quench/lean-burn combustors applied to subsonics. GE has already done this with the Twin Annular Premixing Swirler (TAPS) combustor on the GEnx; see:

http://www.techtransfer.berkeley.edu/aviation05downloads/Dodds.pdf

Quoting Starglider (Thread starter):
Taking total flight time into account, of which cruise is the longest (most efficient) segment, would it be fair to say that less pollution takes place flying an SST long distance for 2 to 3 hours than a lumbering subsonic wide-body for 8 to 9 hours?

In short, no. From the publicly available information on NASA's HSCT program, it appears that the faster cruise speed cannot make up for the lower lift-drag ratio and higher structural weight of the aircraft and the higher specific fuel consumption of a low-bypass turbofan. It appears inevitable that an SST will burn significantly more fuel per seat than a subsonic on an equivalent flight.

Given that, an SST has a negative environmental impact, since emissions are in most cases directly proportional to fuel burn. Compared to a fleet of subsonic jumbos, a fleet of SSTs would reduce the number of aircraft needed to move a given number of people, since each SST is making twice as many flights in a given time period, but the total amount of fuel burned would increase. Add to this the extremely difficult noise problem, which was ultimately NASA's biggest stumbling block on the HSCT, and it becomes clear that environmental compatibility will be a hurdle for a new SST, not a selling point.

--B2707SST

[Edited 2006-12-23 05:47:26]


Keynes is dead and we are living in his long run.
User currently offlineStarlionblue From Greenland, joined Feb 2004, 17055 posts, RR: 67
Reply 5, posted (7 years 9 months 1 week 6 days 9 hours ago) and read 1838 times:

Quoting B2707SST (Reply 4):
Given that, an SST has a negative environmental impact, since emissions are in most cases directly proportional to fuel burn. Compared to a fleet of subsonic jumbos, a fleet of SSTs would reduce the number of aircraft needed to move a given number of people, since each SST is making twice as many flights in a given time period, but the total amount of fuel burned would increase. Add to this the extremely difficult noise problem, which was ultimately NASA's biggest stumbling block on the HSCT, and it becomes clear that environmental compatibility will be a hurdle for a new SST, not a selling point.

Very well explained as usual B2707SST.

I would think that if the noise problem is dealt with, enough people would pay a premium to fly in SS Business Jets and such that it would eventually attract the investment needed to fix the other problems. But that's just a theory. Right now there is not enough investment to really get very far since the amount of money needed is so large. But if someone makes a noise breakthrough the floodgates may well open.



"There are no stupid questions, but there are a lot of inquisitive idiots."
User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 6, posted (7 years 9 months 1 week 4 days 22 hours ago) and read 1775 times:

Quoting B2707SST (Reply 4):
From the publicly available information on NASA's HSCT program, it appears that the faster cruise speed cannot make up for the lower lift-drag ratio and higher structural weight of the aircraft and the higher specific fuel consumption of a low-bypass turbofan. It appears inevitable that an SST will burn significantly more fuel per seat than a subsonic on an equivalent flight.

Thanks for the link regarding TAPS. Interesting development.

Indeed, supersonic aircraft have a lower L/D ratio when compared to subsonic aircraft. But apart from engine technologies there are technological advances in materials such as high temperature composites and aerodynamics that could narrow the gap considerably.

For instance, weight can be reduced by using advanced composites in the airframe. An SST may be less efficient at subsonic and transonic speeds but efficiency (range factor) increases considerably at higher supersonic speeds. Even though L/D ratio is lower, the higher speed compensates for it. So it is important that an SST reaches its designed cruising speed/altitude as soon as possible (the XB-70 was able to achieve this in 25 minutes, timed from moment of rotation).

At supersonic speeds the inlet pressure recovery becomes a very important factor regarding total thrust and specific fuel consumption. Above M2.8 the portions of thrust provided by the inlet, engine and nozzle are about 60, 25 and 15 percent respectively. Virtually coming close to a ramjet.

Another feature that improves range (lowers fuel burn) is compression lift. For instance, when comparing Concorde performance with that of the XB-70 there was a considerable advantage for the latter regarding range factor. Since the aircraft cruise range is proportional to the range factor, a relative drop in efficiency may be offset by a correspondent increase in Mach number.

Range factor= Mach no. x (L/D) in the graph.

Calculations indicate for Subsonic jet transports:14.4, B-52G:16, Concorde:17, XB-70:24. In case of the XB-70, note that L/D was measured at M2.39. The number is even higher at M3 (M3 x (10)=30) See graphs below:

Big version: Width: 600 Height: 589 File size: 35kb

Big version: Width: 800 Height: 269 File size: 38kb



Compression lift also defines the choice of inlet, namely the integrated (2D) wedge shaped inlet instead of (3D) podded inlet design which many designs seem to advocate (apparently for ease of maintenance). Although the integrated inlet is larger/heavier (weight may perhaps be reduced by use of advanced composites), the duct structure serves many purposes. The landing gear are housed in the basic engine nacelle, saving space in the upper fuselage, and wing loads as well as duct loads are carried by the structure. In general the drag of the integrated system can be less and good duct characteristics easier to obtain.

In the XB-70's case the Inlet Air Induction System also contributed in drag reduction by using boundary layer control (BLC) to remove turbulent airflow from the inlet ramp doors. This turbulent air was drawn through the porous ramp doors. The boundary air ducted aft increased the base area pressures and reduced the boat tail drag. This feature alone increased the Lift-Drag ratio by 2 percent at cruising speed.

Empirics of the XB-70 program gives us, in a reverse sort of way, some measure of what can be done to improve inlet efficiency. When this aircraft was changed from a weapon system to a test vehicle, specifications on duct smoothness and on sensitivity of inlet control was relaxed. Correspondingly, performance goals of the XB-70 external-internal inlet, which called for pressure recoveries of 0.93 at takeoff and 0.90 at Mach 3, were rewritten to 0.88 and 0.87, respectively. These apparently small differences in a performance criterion had a large effect on the aircraft design. For the XB-70, 1 percent in presure recovery represents about 1.65 percent in thrust. The 5 percent downgrading of pressure recovery computes to about 5000 lb of penalty in payload. At Mach 3 cruise, better pressure recovery increases payload through lower fuel consumption, e.g. lower emissions. Were the XB-70 powerplant to carry a 3-percent more efficient inlet, about 5000 lb less fuel would be needed to cruise over a typical 3000-4000 mile range.

Now imagine the above mentioned type of aircraft redesigned as a next generation SST with state-of-the-art technologies and engines.

Merry Xmas
Starglider


User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 7, posted (7 years 9 months 1 week 3 days 11 hours ago) and read 1730 times:

Quoting Starglider (Reply 6):
The boundary air ducted aft increased the base area pressures and reduced the boat tail drag. This feature alone increased the Lift-Drag ratio by 2 percent at cruising speed.

In all fairness, i should add that at speeds near Mach 1.2 boat tail base negative pressure coefficients were measured which were 3 times higher than predicted. This means increased boat tail base drag (added to all other drag factors already present in this speed regime), reducing Lift-Drag ratio further at that speed (see graph in reply 6). A factor which can seriously reduce range if acceleration performance in this regime is poor. Its obvious and essential to climb/accelerate through this speed range (of M1 to M1.4) as soon as possible. How soon depends on available transonic excess thrust (inlet throat setting accuracy/engine performance) and aircraft weight. At M1.7 the increase in base pressure started to contribute slightly to range increase. At speeds between M2.6 and M3 the 2 percent additional contribution to L/D ratio was obtained.


User currently offlineB2707SST From United States of America, joined Apr 2003, 1369 posts, RR: 59
Reply 8, posted (7 years 9 months 1 week 2 days 21 hours ago) and read 1704 times:

Quoting Starglider (Reply 6):
An SST may be less efficient at subsonic and transonic speeds but efficiency (range factor) increases considerably at higher supersonic speeds. Even though L/D ratio is lower, the higher speed compensates for it. So it is important that an SST reaches its designed cruising speed/altitude as soon as possible (the XB-70 was able to achieve this in 25 minutes, timed from moment of rotation).

It is true that M*(L/D) for an SST is (or should be) greater than for a subsonic. For instance, the 2707-100's M*(L/D) was 2.7*(8.5) = 23 compared to the 747's 0.85*(18) = 15.3 or the 2707's own 0.85*(16) = 13.6 at subsonic cruise speed of Mach 0.85.

There is some hope on the aerodynamic front: modern SST designs seem to have L/Ds in the 10-11 range instead of the 8-9 of first-generation SSTs. British Aerospace had reached this point in the mid-1990s, as shown below:



Apparently, as had NASA in its HSCT studies:



Lockheed also cited a figure of 10-11 as the theoretical maximum for a delta-winged SST configuration, so this range seems very reasonable for a new SST. However, cruise speed targets have also come down since the 1960s, when Mach 2.7-3.0 was the goal for the US/SST program. More recently, independent evaluations of the HSCT warned that even Mach 2.4 would be a challenge and that Mach 2.2 might be a wiser choice in terms of cost/benefit, as the flight time difference is relatively minor compared to the reduction in stagnation temperature. Thus, M*(L/D) of 2.4*(11) = 26.4 seems to be about the upper bound, with 2.2*(10.5) = 23 probably a more reasonable estimate.

So we can say that, all else equal, SSTs outperform subsonics on a M*(L/D) basis, at least in cruise. But all else is not equal: low- or zero-bypass supersonic engines have higher SFCs than their subsonic counterparts, and supersonic airframes are significantly heavier. TSFC for the GE4, developed from the XB-70's YJ93, was 1.04 at Mach 2.7 cruise, with an axisymmetric inlet system achieving 91% pressure recovery (higher values were possible but were sacrificed for greater unstart stability). Concorde's Olympus 593 Mk. 610 achieved 1.18. These compare fairly well with the contemporary JT3D's 0.85 but do not look so good against today's GEnx at around 0.535.

I am confident GE, RR, or P&W could beat the Olympus or the GE4 if they built another straight turbojet with modern technology. The problem is that a new straight turbojet will not be acceptable environmentally (noise, emissions) and economically (subsonic fuel burn). Some amount of bypass flow will be necessary, and this will necessarily pull down efficiency at cruise. According to knowledgable A.net poster and former Concorde engineer GDB, RR does not expect to appreciably beat the Olympus' efficiency at supersonic cruise despite all our technological advances since the 1960s, since the design compromises necessary today claw back most, if not all, of those gains.

Weight is also a major issue. High-temp composites, though currently unproven, could offer significant gains over titanium, but a new SST will still be substantially heavier than an equivalent subsonic. The greater temperatures and stresses of high-speed flight are certain to incur a weight penalty, and the higher SFC of SST engines will require a heavier airframe to carry the larger fuel load. The very solutions to problems like noise and engine efficiency -- leading and trailing high-lift devices, variable-bypass inlets, mixer-ejector silencing nozzles -- will add more weight and complexity. Similarly, overpowering an SST in order to rocket to cruise altitude would require heavy and fuel-thirsty engines and might reduce overall efficiency compared to powerplants more closely matched to cruise thrust requirements.

Given all these challenges, and especially given that the bar is moving rising faster than progress is being made (e.g. subsonic airliners seem to be growing more efficient more quickly than SST proposals), I remain pessimistic on the prospects for a new SST. The introduction of the 787 alone will require a technological leap on the part of its competitors, both subsonic and supersonic. Moreover, the factors above are all technical: I haven't even touched on the immense expense of such a project and the limited and uncertain market for premium-priced supersonic travel. Put simply, the 787 and A350 are far wiser investments than an SST program costing four times as much with potential unit sales in the hundreds rather than thousands.

I would not be at all surprised to see an SSBJ or two in the next decade to skim off the very top end of the market; a project of that scale and scope is far more rational than a new SST would be. Barring a stunning and unforeseen breakthrough in propulsion technology, laminar flow, materials, etc., a full-sized supersonic airliner is at least several decades away.

--B2707SST



Keynes is dead and we are living in his long run.
User currently offlineLehpron From United States of America, joined Jul 2001, 7028 posts, RR: 21
Reply 9, posted (7 years 9 months 1 week 2 days 21 hours ago) and read 1700 times:

I've always wondered, what does the issue with NOx and engines has to do with? I mean if is it the temperature of exhaust of flying within the ozone layer? Would less temperate engine exhaust and/or flying above or below the Ozone layer help? Or is it a product of combustion issue including temperature and ozone?


The meaning of life is curiosity; we were put on this planet to explore opportunities.
User currently offlineAvt007 From Canada, joined Jul 2000, 2132 posts, RR: 5
Reply 10, posted (7 years 9 months 1 week 2 days 20 hours ago) and read 1699 times:

I would be more than happy to fly an SST (if my boss would approve my expense claims). The technology exists, and I think that the enviromental concerns, both noise and pollution , will be beat. But even though we already have the technology, it isn't economically feasable at this point. But if you guys figure out a way, sign me up!

User currently offlineB2707SST From United States of America, joined Apr 2003, 1369 posts, RR: 59
Reply 11, posted (7 years 9 months 1 week 2 days 19 hours ago) and read 1695 times:

Quoting Lehpron (Reply 9):
I've always wondered, what does the issue with NOx and engines has to do with? I mean if is it the temperature of exhaust of flying within the ozone layer? Would less temperate engine exhaust and/or flying above or below the Ozone layer help? Or is it a product of combustion issue including temperature and ozone?

Nitrogen oxides (NO or NO2) are produced in the combustion chamber when molecular nitrogen (N2) and oxygen (O2) combine at very high temperatures. When these products are released into the atmosphere, they catalytically destroy ozone (O3) molecules that shield the surface from harmful UV rays:

NO2 + O3 -> 2 O2 + NO
NO + O3 -> O2 + NO2

Since the nitrogen oxides are not consumed in the reaction and just cycle back and forth between nitic oxide and nitrogen dioxide, they can destroy large numbers of ozone molecules before being purged from the stratosphere.

NOx formation requires very high temperatures and occurs when fuel and air combine in the perfect "stoichiometric" ratio, which designers would typically favor for maximum efficiency. The solution is to burn the fuel either rich or lean, sometimes in stages, to keep the flame temperature down.

The two primary combustor designs proposed for NOx control are lean premixed prevaporized, which (like it sounds) vaporizes the fuel without igniting it, then mixes it with more than enough air before ignition; and rich-burn/quick-quench/lean-burn, which allows fuel to burn with insufficient air before snuffing out the combustion with a cold air blast, then allowing combustion to resume with a lean mixture. Both of these approaches reduce the amount of stoichiometric (i.e. perfect fuel-air) combustion and prevent the temperature spikes that create NOx.

--B2707SST



Keynes is dead and we are living in his long run.
User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 12, posted (7 years 9 months 1 week 1 day 8 hours ago) and read 1657 times:

Quoting B2707SST (Reply 8):
I am confident GE, RR, or P&W could beat the Olympus or the GE4 if they built another straight turbojet with modern technology. The problem is that a new straight turbojet will not be acceptable environmentally (noise, emissions) and economically (subsonic fuel burn). Some amount of bypass flow will be necessary, and this will necessarily pull down efficiency at cruise. According to knowledgable A.net poster and former Concorde engineer GDB, RR does not expect to appreciably beat the Olympus' efficiency at supersonic cruise despite all our technological advances since the 1960s, since the design compromises necessary today claw back most, if not all, of those gains.

Thanks for your detailed reply. I agree there are still many issues to be solved. However, recently progress has been made in powerplant studies:

Quoting NASA report TM-2004-213139:
"A parametric study is conducted to evaluate a mixed-flow turbofan equipped with a Supersonic Through-Flow Rotor and a Supersonic Counter-Rotating Diffuser (SSTR/SSCRD) for a Mach 2.4 civil transport. Engine cycle, weight, and mission analyses are performed to obtain a minimum takeoff gross weight aircraft. With the presence of SSTR/SSCRD, the inlet can be shortened to provide better pressure recovery. For the same engine airflow, the inlet, nacelle, and pylon weights are estimated to be 73 percent lighter than those of a conventional inlet. The fan weight is 31 percent heavier, but overall the installed engine pod weight is 11 percent lighter than the current high-speed civil transport baseline conventional mixed-flow turbofan. The installed specific fuel consumption of the supersonic fan engine is 2 percent higher than that of the baseline turbofan at supersonic cruise. Finally, the optimum SSTR/SSCRD airplane meets the FAR36 Stage 3 noise limit and is within 7 percent of the baseline turbofan airplane takeoff gross weight over a 5000-n mi mission."

Below are some comparisons between several supersonic engine configurations and SSTR/SSCRD:



Big version: Width: 600 Height: 458 File size: 66kb

Big version: Width: 791 Height: 399 File size: 66kb


Below are comparisons between a MFTF (Mixed-Flow Turbo Fan) and a SSTR/SSCRD engine:

Big version: Width: 367 Height: 424 File size: 25kb
Big version: Width: 551 Height: 338 File size: 35kb
Big version: Width: 653 Height: 469 File size: 65kb


In the SSTR/SSCRD section of the engine there are three blades rows:
- a supersonic through-flow rotor;
- a supersonic counter rotating diffuser and,
- a stator.

The diffuser rotates in the opposite direction of the supersonic rotor. The supersonic rotor (SSTR) can operate with subsonic or supersonic inflow. For subsonic inflow, the flow is accelerated to supersonic conditions at the SSTR exit. At high rotational speeds, the SSTR fixed the engine corrected airflow to nearly constant value. For supersonic inflow, the engine corrected airflow can vary but the rotor face Mach number must be maintained at 1.35 or above to avoid fan unstart. The rotor exit outflow usually is supersonic except at very low part power operation. The supersonic rotating diffuser (SSCRD) is started with supersonic inflow and diffuses the flow to subsonic absolute Mach numbers. The exit guide vanes then renove the swirl from the subsonic flow before it enters the high-pressure compressor and bypass duct.

Comparing conventional MFTF with SSTR/SSCRD:
From sea level static to Mach 0.4, the pressure recovery was significantly lower for the SSTR/SSCRD inlet as a result of large lip loss. The pressure recovery improved from Mach 0.4 to low supersonic Mach numbers because of the shorter inlet design and a smaller diffuser loss but fell off between Mach 1.4 to 1.75 because unstarted external shock losses began to rise. The SSTR/SSCRD inlet has a starting Mach number of M 1.75. At Mach 1.9, the rotor starts and the shock wave moves into the fan and disappears. From M 1.75 to 2.4, the recovery was higher because of a single reflected internal obligue shock wave as opposed to the multiple shocks on the MFTF inlet. Bypass doors were opened between Mach 0.9 to 2.1 to lower drag and to assist in matching the supplied airflow to the engine demand.

Drag coefficient for the SSTR/SSCRD inlet is four times greater than that for the conventional axisymetric inlet in the transonic region. This is attributed to a fixed-geometry inlet design, which requires that a large amount of air be spilled by the inlet. For the same capture area, the SSTR/SSCRD engine receives less airflow than that of the MFTF.

Between Mach 2.1 and 2.4, the drag coefficient is minimal because little air is being spilled or bypassed. A benefit of the SSTR/SSCRD inlet is that it requires no inlet bleed air for stability anywhere in the flight regime. For subsonic inflow, the Mach number at the rotor face is constant at Mach 0.8, which results in constant engine corrected airflow from takeoff until the fan starts. Once the fan starts, supersonic inflow occurs and all the air captured by the inlet goes through the fan.

There were 9 different SSTR/SSCRD engine variants studied of which SSTR/SSCRD7 has the right combination of engine pod weight and total fuel weight to provide the optimum airplane with the lowest TOGW (take-off gross weight) at sea level static bypass ratio of 0.56 and a FPR (fan overall pressure ratio) of 3.05. The SSTR/SSCRD7 also has the highest overall efficiency of the 9 variants studied and this has an effect on the aircraft TOGW. The cruise thrust specific fuel cosumption of the SSTR/SSCRD7 airplane is 2 percent higher than that of the baseline turbofan. A further reduction in FPR cannot be accomplished because of the SSTR/SSCRD fan operation limits. The baseline conventional MFTF has a lower TOGW at a sea level static bypass ratio of 0.80 (see graphs above).

The mission sized SSTR/SSCRD7 propulsion weight is 33 percent lower, resulting in an overall 3 percent lower operating empty weight than that of the MFTF. The SSTR/SSCRD7 airplane also has 12 percent better climb fuel because the airplane has more power to climb at a faster rate. However, subsonic and supersonic cruise fuel consumption are 15 percent and 18 percent higher, respectively. Supersonic cruise and reserve fuel weights make up a substantial 36 percent of the SSTR/SSCRD7 airplane TOGW.

Overall, SSTR/SSCRD7 engine performance shows promising results compared with the conventional MFTF in terms of a comparable airplane TOGW while still satisfying the FAR36 Stage 3 noise requirement at take-off. NOx formation can be reduced by decreasing the burner temperature, which can be accomplished by lowering the temperature of the gas coming into the burner and by operating at a leaner fuel-to-air ratio.

A take-off door on the supersonic inlet of the SSTR/SSCRD7 was used to increase the take-off thrust by providing better pressure recoveries from sea level static to Mach 0.4 with the same ratio of engine airflow to inlet capture area. The result was a smaller inlet capture area and a lower sea level static engine corrected airflow accompanied by an increase in net thrust and decrease in TCFC at take-off. The installed fuel flow at supersonic cruise was 6 percent lower, which resulted in a reduction of 4 percent in airplane TOGW. Therefore, the take-off door augmented take-off thrust, improved the supersonic cruise fuel consumption, and more important, lowered the airplane TOGW.

The supersonic cruise aircraft used for the 5000 n mi range, Mach 2.4 at 60.000 ft study was a NASA variant of the HSCT that carries 301 passengers.

Perhaps a study could be conducted to strap the SSTR/SSCRD7 to an SST concept using compression lift, perhaps increasing total aerodynamic efficiency even further?

Regards,
Starglider


User currently offlineB2707SST From United States of America, joined Apr 2003, 1369 posts, RR: 59
Reply 13, posted (7 years 9 months 1 week 1 day 3 hours ago) and read 1641 times:

Quoting Starglider (Reply 12):
The mission sized SSTR/SSCRD7 propulsion weight is 33 percent lower, resulting in an overall 3 percent lower operating empty weight than that of the MFTF. The SSTR/SSCRD7 airplane also has 12 percent better climb fuel because the airplane has more power to climb at a faster rate. However, subsonic and supersonic cruise fuel consumption are 15 percent and 18 percent higher, respectively. Supersonic cruise and reserve fuel weights make up a substantial 36 percent of the SSTR/SSCRD7 airplane TOGW.

This is a very interesting study, but based on the figures cited above, the "conventional" mixed-flow turbofan with variable-position axisymmetric inlet appears to be a superior option. The SSTR/SSCRD results in 15-18% increases in cruise mode SFC values, which more than offset its lower climb fuel burn and lower OEW, as shown by the nearly 20% increase in block fuel and 7.5% increase in MTOW for an identical payload/range. The final HSCT design, the Technology Concept Aircraft, specified a 740,000-lb. MTOW for 300 pax and 5,000 nmi range, suggesting some improvement from the MTF design described above.

According to NASA documentation, the HSCT MTF/mixer-ejector combination could meet Stage III noise requirements at this gross weight. NASA could not, however, achieve Stage IV compliance, which will in effect when a new SST enters service (and even more stringent standards may still be introduced). This was one of the key drivers in the decision to terminate the HSR program.

--B2707SST



Keynes is dead and we are living in his long run.
User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 14, posted (7 years 9 months 6 days 12 hours ago) and read 1602 times:

Quoting B2707SST (Reply 13):
This is a very interesting study, but based on the figures cited above, the "conventional" mixed-flow turbofan with variable-position axisymmetric inlet appears to be a superior option.

It may appear so but apart from the comparatively higher TSFC of the SSTR/SSCRD7 it does have advantages over the "conventional" mixed flow turbofan and its inlet design. Apart from lower installed weight:

- A stable inlet through the entire flight regime. The fixed inlet has a simpler design when compared to the axisymmetric, multi-shock spike inlet which is complex and its performance is good only over a very limited range of angles of attack.

In the graph below ratio of engine airflow to inlet capture area schedule compare closely:

Big version: Width: 722 Height: 484 File size: 38kb


- A stable inlet is a positive contribution to flight safety and operation. It means less chance of an inlet "unstart" which is comfortable to know for crew and passengers. An XB-70 pilot once described an inlet "unstart" as "being T-boned by a locomotive and thrown off a cliff".

The graph below shows primary inlet fan Mach number through the entire flight regime:

Big version: Width: 717 Height: 503 File size: 35kb


The Mach number at the rotor face is constant at M 0.8 until the fan "starts".

- Primary inlet pressure recovery is more efficient at transonic and cruise speed than for the "conventional" inlet, although it falls off between M 1.4 to 1.75 because unstarted external shock losses began to rise. The inlet "starts" at M 1.75. At M 1.9 the rotor starts and the shock wave moves into the fan and disappears. From M 1.75 to 2.4, the pressure recovery is higher because of a single reflected oblique shock as opposed to multiple shocks in the MFTF inlet:

Big version: Width: 748 Height: 500 File size: 44kb


- NOx formation can be reduced by decreasing the burner temperature, which can be accomplished by lowering the temperature of the gas coming into the burner and by operating at a leaner fuel-to-air ratio.

- adding a takeoff door to the SSTR/SSCRD7 improved takeoff performance with a lower TSFC at take-off. Fuel flow at supersonic cruise was reduced by 6 percent and TOGW by 4 percent when compared to the initial figures shown in reply 12.

The engine cycle performance for the SSTR/SSCRD was calculated using the NASA Engine Performance Program (NEPP) and the SSTR/SSCRD components were estimated using the engine weight code WATE-2. All info is based on calculations and as far as i know, no such engine was ever built. Perhaps development has moved on (proprietary data?) and who knows, performance may have improved?


Starglider


User currently offlineBlackbird From , joined Dec 1969, posts, RR:
Reply 15, posted (7 years 9 months 6 days 5 hours ago) and read 1588 times:

Just out of curiousity, do afterburners do produce more NOx than a regular combustor?

P.S. I know this sounds stupid, but could a TAPS type system be fitted to an afterburner (purely hypothetically speaking) paired in a series of rings, etc.

Andrea K


User currently offlineStarglider From Netherlands, joined Sep 2006, 678 posts, RR: 44
Reply 16, posted (7 years 9 months 3 days 23 hours ago) and read 1552 times:

Quoting Blackbird (Reply 15):
Just out of curiousity, do afterburners do produce more NOx than a regular combustor?

Aircraft fitted with afterburner systems for increased thrust have been observed to have NOx emissions with a higher proportion of nitrogen dioxide (NO[2]) than non-augmented aircraft. These emissions are generally characterised by a brown plume.

Quoting Blackbird (Reply 15):
P.S. I know this sounds stupid, but could a TAPS type system be fitted to an afterburner (purely hypothetically speaking) paired in a series of rings, etc.

Good question, but if it would work? Perhaps it depends on the speeds/altitudes the afterburner is designed to operate.

TAPS is a main combustor design in the core of the engine and drives the turbine(s). Mixing of air and fuel in those combustors use highly compressed air. The afterburner receives gases at at relatively low temperature and pressure since a certain amount of heat (reducing gas pressure) has been converted into energy to drive the turbine(s).

Since the exhaust gas already has reduced oxygen due to previous combustion, and since the fuel is not burning in a highly compressed air column, the afterburner is generally inefficient compared with the main combustor. Afterburner efficiency also declines significantly if, as is usually the case, the tailpipe pressure decreases with increasing altitude. In that respect, i doubt TAPS would have any useful benefit.

However, as a counter-example the SR-71 had reasonable efficiency at high altitude in after-burning mode ("wet") due to its high speed (mach 3.2) and hence high pressure due to ram effect.

In most cases it is questionable if TAPS could be successfully installed as part of the afterburner. TAPS replaces existing combustors in the core engine (little or no weight impact). TAPS would be an addition to the afterburner, adding complexity and increasing its weight. Furthermore, TAPS is designed to lower combustion temperature (lowering NOx). In case of the afterburner it would be counter-productive to do so, for it reduces the performance of the afterburner.

Hypothetically, if it could be installed, then most likely for high speeds/altitudes (SR-71 type aircraft).

Taken from an article to reflect on my "doodling":
"Unlike the main combustor, where the integrity of the downstream turbine blades must be preserved, an afterburner can operate at the ideal maximum (stoichiometric) temperature (i.e. about 2100K(3780R)). Now, at a fixed total applied fuel:air ratio, the total fuel flow for a given fan airflow will be the same, regardless of the dry specific thrust of the engine. However, a high specific thrust turbofan will, by definition, have a higher nozzle pressure ratio, resulting in a higher afterburning net thrust and, therefore, a lower afterburning specific fuel consumption. However, high specific thrust engines have a high dry SFC. The situation is reversed for a medium specific thrust afterburning turbofan: i.e. poor afterburning SFC/good dry SFC. The former engine is suitable for a combat aircraft which must remain in afterburning combat for a fairly long period, but only has to fight fairly close to the airfield (i.e cross border skirmishes) The latter engine is better for an aircraft that has to fly some distance, or loiter for a long time, before going into combat. However, the pilot can only afford to stay in afterburning for a short period, before his/her fuel reserves become dangerously low."

Although it is off topic in relation to your question, there is probably more benefit in a TAPS equipped supercruise, turbofan engine to lower NOx emissions as much as possible. Perhaps removing the need for an afterburner all together. At least for aircraft cruising between Mach 1.6 and 2.4.

Best wishes for 2007 and beyond  Smile
Starglider

[Edited 2007-01-02 03:31:34]

User currently offlineGDB From United Kingdom, joined May 2001, 13220 posts, RR: 77
Reply 17, posted (7 years 9 months 3 days 7 hours ago) and read 1524 times:

Well I cannot add anything to the excellent, highly detailed, replies so far.

Other than something anecdotal.
Two or three years ago, how much attention was given, in the serious media, to airliner emissions, compared to now?
My unscientific view is that this issue is much higher up the agenda now, within the past year/18 months it has become much higher profile, it will provoke a political reaction.
Often pointed out is that despite the very significant gains in lower emissions of modern airliners, this is totally reversed due to the sheer growth of air travel.

Try selling a new SST programme in this, pardon the pun, enviroment.


Top Of Page
Forum Index

Reply To This Topic SST Propulsion Emissions And Noise.
Username:
No username? Sign up now!
Password: 


Forgot Password? Be reminded.
Remember me on this computer (uses cookies)
  • Tech/Ops related posts only!
  • Not Tech/Ops related? Use the other forums
  • No adverts of any kind. This includes web pages.
  • No hostile language or criticizing of others.
  • Do not post copyright protected material.
  • Use relevant and describing topics.
  • Check if your post already been discussed.
  • Check your spelling!
  • DETAILED RULES
Add Images Add SmiliesPosting Help

Please check your spelling (press "Check Spelling" above)


Similar topics:More similar topics...
Sudden Interest In SST And RAM Jets posted Fri Jun 23 2006 04:44:41 by Bhill
Noise "Stages" And Rules posted Thu Sep 25 2003 20:38:49 by Bigphilnyc
Canadian Noise Regulations And The AN-124 posted Sun Aug 5 2001 19:17:29 by LY744
737 Flap Settings And Company Policy posted Tue Dec 19 2006 20:09:22 by Futureatp
Doctrine, Procedures And Airmanship posted Tue Dec 19 2006 14:13:56 by Pihero
Cabin Lighting At Take-off And Landing At Night posted Mon Dec 18 2006 03:04:40 by Goodday
UDFs And Type Ratings posted Sun Dec 17 2006 22:35:25 by 2H4
CHQ Operation And CAT II posted Fri Dec 15 2006 13:44:30 by IAHFLYR
What's Different Between BBJ And 737-700ER? posted Sat Dec 9 2006 18:44:52 by DfwRevolution
Electric Brakes And Deadstick Landings posted Fri Dec 8 2006 06:47:44 by WingedMigrator

Sponsor Message:
Printer friendly format